NASA TECHNICAL NOTE \ NASA TN D-6078

INVESTIGATION OF A MIXED-COMPRESSION AXISYMMETRIC INLET SYSTEM AT MACH NUMBERS 0.6 TO 3.5

. . ' / .

i ! N A T I O N A L A E R O N A U T I C S A N D S P A C E A D M I N I S T R A T I O N W A S H I N G T O N , D. C. NOVEMBER 1970 J

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TECH LIBRARY WFB, NM

1. Report No. I 2. Government Accession No. NASA TN D-6078 I

4. -Title and Subtitle 5. Report Date

INVESTIGATION O F A MIXED-COMl'RESSlON AXISYMMETRIC INLET November 1970 SYSTEM AT MACH NUMBERS 0.6 TO 3.5 I 6. Performing Organization Code

7. Author(s)

Donald 8. Smeltzer and Norman E. Sorensen 8. Performing Organization Report No. I A-3516

10. Work Unit No. 9. Performing Organization Name and Address 720-03-01-10-00-21

NASA Ames Research Center Moffett Field, Calif., 94035

I 11. Contract or Grant No.

13. Type of Report qnd Period Covered 12. Sponsoring Agency Name and Address Technical Note

National Aeronautics and Space Administration Washington, D. C., 20546

I 14. Sponsoring Agency Code

" I 15. Supplementary Notes

16. Abstract . . ..

A 20-inch (50.8 cm) capture diameter model of a mixed-compression axisymmetric inlet systcm has been tested. The design Mach number was 3.5 and off-design performance was obtained by translation of the cowl. Thc inlet system was 1.4 capture diameters long, measured from the cowl lip to the engine face. Vortex generators were installed just downstream of the throat to reduce the total pressure distortion at the engine face. A boundary-layer removal system was provided on both the cowl and centerbody surfaces of the subsonic and supersonic diffuser. An engine airflow bypass system was located on the cowl surface just upstream of the engine face.

The major performance parameters of bleed mass-flow ratio, total-pressure recovery, and total-pressure distortion were determined as a function of bypass mass-flow ratio. Other results obtained were transonic additive drag, inlet tolerance to change in angle of attack, boundary-layer profiles,and surface-pressure distributions. At Mach number 3.5 and zero bypass, total-pressure recovery ranged from 84.5 to 89.5 percent at the engine face with a bleed mass-flow ratio of 14 to 21 percent. The total-pressure distortion level was about 5 percent. Testing was conducted at a tunnel total pressure of 15 psia which, at Mach number 3.5, corresponded to a unit Reynolds number of about 1.6X106/ft. 111 the supersonic range, data wcre obtained at Mach number incrcments of 0.15 between Mach numbers 1.5 and 3.5, and i n the transonic range, at Mach number increments of 0.1 between Mach numbers 0.6 and 1.3. At all Mach numbers, data were obtained at four angles of attack between 0' and 8"

;7."Key Words (Suggested by Author(s) ) ..

Air breathing Ducts Inlets Boundary layers Propulsion Internal flow

18. Distribution Statement

Unclassified - Unlimited

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SYMBOLS

m

P

Pt

K

- r R

- X R

capture area

minimum area

local duct area normal to the inlet centerline

additive drag coefficient based on Ac

capture diameter 20 in. (50.8 cm)

local diameter

local height

local dimension

Mach number

mass flow

static pressure

total pressure

- 't2max total pressure distortion parameter, -

Pt 2

capture area radius

ratio of local radius to capture area radius

ratio of axial distance from the tip of the centerbody to the capture-area radius

ratio of axial distance from the cowl lip to the capture-area radius

axial distance from the cone tip to the cowl lip ratioed to the capture-area radius

incremental, X

iii

a

auns

( 7

angle of attack, deg

angle of attack for incipient unstart, deg

average value

Subscripts

bl bleed

bP bypass

d downstream

i inlet lip (measured)(transonic results only)

1 local

0 inlet lip (theoretical)

U upstream

1 throat

2 engine face

00 free stream

NOTE: The letters A, B, and C on the plotted and tabulated data refer t o progressively restricted bleed exit settings, A being the maximum flow condition and C the most restricted.

iv

INVESTIGATION OF A MIXED-COMPRESSION AXISYMMETRIC INLET SYSTEM

AT MACH NUMBERS 0.6 TO 3.5

Donald B. Smeltzer and Norman E. Sorensen

Ames Research Center

SUMMARY

A 20-inch (50.8 cm) capture diameter model of a mixed-compression axisymmetric inlet system was tested at the design Mach number of 3.5 and off-design conditions. The variable geometry inlet system was 1.4 capture diameters long measured from the cowl lip to the engine face. Vortex generators were used just downstream of the throat to reduce the total-pressure distortion at the engine face. A boundary-layer removal system was provided on both the cowl and centerbody surfaces of the subsonic and supersonic diffuser. An engine airflow bypass system was located just upstream of the engine face. Struts similar to those used on a full-scale inlet system upstream of the engine face were not provided.

The main test objective was to investigate the major inlet parameters of bleed mass flow, total-pressure recovery, and total-pressure distortion at the engine face with various amounts of bypass mass flow. Tests were conducted over the Mach number range 0.6 to 3.5 at a wind-tunnel pressure of 15 psia which, at Mach number 3.5, corresponded to a unit Reynolds number of about 1 .6X106/ft. Results were obtained at angles of attack from 0" and 8".

The supersonic diffuser was designed with the aid of a computer program which employs the method of characteristics. The subsonic diffuser was designed to have a linear variation of Mach number with distance from the end of the throat to the engine face.

At Mach number 3.5 and with zero bypass the maximum total-pressure recovery ranged frorn 84.5 to 89.5 percent as the bleed mass-flow ratio increased from 14 to 21.5 percent and the total-pressure distortion remained approximately constant at 5 percent. At Mach numbers less than 3.5 and with zero bypass, maximum total-pressure recovery was generally about the same or slightly better, but total-pressure distortion was generally higher. Total-pressure recovery generally decreased slowly with increasing bypass throughout the Mach number range while total-pressure distortion remained about the same or slightly lower. Results obtained throughout the Mach number range at 2" angle of attack showed only small decreases in maximum total-pressure recovery and about the same total-pressure distortion. However, at larger angles of attack (5" and So), considerable decreases in maximum total-pressure recovery and increases in total-pressure distortion occurred.

Other results indicated that when operated at maximum pressure recovery (contraction ratio 7.22) at Mach number 3.5, the inlet unstarted at approximately 1" angle of attack. A small decrease in contraction ratio and withdrawal of the terminal shock wave downstream permitted operation without unstart to about 2.4" angle of attack with only a small performance penalty.

INTRODUCTION

As the design Mach number of cruise vehicles increases, the achievement of a satisfactory inlet system becomes more difficult. The attainment of high total-pressure recovery and low total-pressure distortion with small amounts of boundary-layer bleed has been demonstrated for inlets designed for Mach numbers 2.5 and 3.0. These results are shown in references I , 2, and 3. However, a t higher design Mach numbers, the requirement that the inlet system be as short as possible results in increasingly severe adverse pressure gradients acting on the boundary layer. Because of this, i t was uncertain as to whether or not high total-pressure recovery and low total-pressure distortion could be achieved without excessive boundary-layer bleed. The main purpose of the study was, therefore, t o investigate the major inlet parameters of total-pressure recovery and total-pressure distortion as a function of boundary-layer bleed. In addition, because an engine airflow bypass system was included in the design, it was desired to investigate the effect of bypass mass flow on the principle performance parameters. Reference 4 compares some of the present results with the results of references 1, 2, and 3.

The supersonic diffuser was designed with the aid of the computer program, which employs the method of characteristics (ref. 5 ) . The subsonic diffuser was designed to yield a linear variation of Mach number with distance. The resulting model had a mixed-compression supersonic diffuser with a 20-inch (50.8 cm) capture diameter and was matched to a very short subsonic diffuser producing a length of 1.4 capture diameters measured from the cowl lip to the engine face. The model was designed so that, at Mach number 3.5, the shock wave from the conical centerbody was just upstream of the cowl lip. The area distributions required for operation throughout the Mach number range were achieved by translation of the cowl. Vortex generators were incorporated on the cowl and centerbody surfaces just downstream of the throat to reduce the eotal-pressure distortion at the engine face. A boundary-layer removal system to control boundary-layer separation was provided on both the cowl and centerbody surfaces. A bypass airflow system for simulating engine airflow matching and inlet restarting was provided on the cowl surface just upstream of the engine face. The model is shown in figure 1 installed in one of the supersonic wind tunnels.

At Mach number 3.5 and 0" angle of attack, various combinations of bleed hole patterns were investigated. This involved varying the expanse and location of the bleed holes in both the supersonic diffuser and the throat. Combinations were investigated until i t became apparent that it would be difficult to further improve performance. Data were recorded at off-design Mach numbers and at angle of attack only with the bleed hole pattern that gave the best performance at Mach number 3.5 and 0" angle of attack and all reported results are for this single bleed hole pattern. Measurements were made of total-pressure recovery and total-pressure distortion at the engine face as a function of bleed mass flow or mass flow at the engine face. The effect of bypass mass flow on these parameters was also determined. Boundary-layer profiles, surface-pressure distributions, and bleed and bypass plenum-chamber pressures were also measured. Inlet sensitivity to unstarting caused by changes in angle of attack was also determined. In addition, experimental determination of transonic additive drag was made.

2

MODEL

Sketches of the model and instrumentation are shown in figures 2(a) and 2(b). The model had a 20-inch (50.8 cm) capture diameter and the required off-design inlet area variations were accomplished by translation of the cowl. The cowl had a sharp 15" lip and could be translated about 0.85 capture diameters. At the exit station a translating sleeve and fixed plug controlled the terminal shock-wave position. The outer shell was attached to four hollow struts mounted on the centerbody sting support. The struts supported the cowl and provided ducting for the centerbody bleed airflow to the free stream. Four separate bIeed zones and a compartmented bypass zone each had a remotely controlled exit which regulated the flow from zero to full flow. Separation of the bleed zones and compartmentation of the bypass zone prevented recirculation of the flow from the higher to the lower pressure regions. To ensure low back pressures at the bleed and bypass exits, fairings were provided as shown in figure 2(a). Further details of the design and test instrumentation are discussed in the following sections.

DESIGN

The prime objective was the attainment of high total-pressure recovery and low total-pressure distortion at the engine face with a minimum amount of bleed mass flow at the design Mach number of 3.5. The performance at lower Mach numbers, however, was not overlooked. Other objectives were to attain low cowl drag and low transonic additive drag and t o keep the inlet system as short as possible.

The design principles used were those that were successful for the inlets of references 1 to 4. These included the following: The supersonic diffuser would be of a mixed-compression type; the initial cone and external cowl angles would be small; and the pressure rise across shock-wave impingement points would be kept low. In addition, flow separation could be controlled by the proper location of boundary-layer bleed, and flow distortion i n the subsonic diffuser could be reduced by vortex generators. The inlet coordinates are presented in table I and the important aspects of the design are considered in the following paragraphs.

Supersonic Diffuser

This portion of the inlet was designed with the aid of the computer program described in reference 5, which employs the method of characteristics. Figure 3 shows the flow field output from the computer program for the design Mach number of 3.5. To achieve low cowl drag an initial internalcowl angle of 0" was selected. The requirement for low spillage drag was satisfied by the selection of a cone with a 10" half angle for the initial surface of the centerbody, x/R = 0 to 1.636. Between x/R = 1.636 and 3.800 a linear rate of change of surface slope with distance was used so that the total turning of the surface was 15" at station x/R = 3.800. The remaining contours of the cowl and centerbody were tested using the computer program until the desired conditions were attained across the inlet throat. These conditions were a uniform Mach number of about 1.25 with essentially parallel flow and a total-pressure recovery of about 0.985. Another design constraint was that the pressure rise across the shock-wave reflections on the centerbody and cowl could not

exceed the value for incipient boundary-layer separation as defined in' reference 6. (Although the data in ref. 6 were obtained for two-dimensional flow, they have been used with apparent success for previous axisymmetric inlet designs.) The resulting design gave a capture mass flow at Mach number 1.0 of 33.5 percent and a cowl translation distance of about 0.67 capture diameter for operation throughout the Mach number range. No boundary-layer compensation was included in the design of the supersonic diffuser since previous experience (refs. 1 , 2, and 3) indicated that with the boundary-layer removal system none would be required.

Subsonic Diffuser

The Mach number at the beginning of the subsonic diffuser was determined from the output of the computer program (fig. 3). In the throat region (x/R = 4.225 to 4.399, the centerbody and cowl surfaces diverged from one another at 2O. Because the cowl surface was a straight line from (x/R), = 1.175 to 1.535, the 2" divergence in the throat was maintained over a range of cowl translation distance. This arrangement located the minimum throat area on the centerbody at about x/R = 4.26 throughout the cowl translation from (x/R)lip = 2.825 to 3.085. Translating the cowl farther aft shifted the minimum throat area to a forward limit of x/R = 4.18. These characteristics are illustrated in figure 4 by the inlet area distributions shown for the various Mach numbers tested. The remainder of the subsonic diffuser, from the aft end of the throat (x/R = 4.395) to the engine face (x/R = 5.650), was designed to have a linear Mach number variation, which was maintained to some degree at off-design Mach numbers. The location (x/R = 5.650) and area of the engine face were dictated by the fact that the same model used for previous tests at Mach number 3.0 was to be modified for the present investigation. Although this resulted in a short subsonic diffuser with an equivalent conical angle of 28" from the beginning of the throat to the engine face, i t was believed that the vortex generators would reduce the total-pressure distortion at the engine face to an acceptable level. The area at the engine face was consistent with current estimates of an engine designed to operate at Mach number 3.5. The resulting contour provided a range of engine-face Mach numbers from about 0.15 at a free stream Mach number of 3.5 to about 0.40 at a free stream Mach number of 1.0. As in the case of the supersonic diffuser, no boundary-layer compensation was included in the design of the subsonic diffuser contour.

Bleed System

Figure 2(b) shows the bleed pattern used for all reported results and figure 3 shows the location of the bleed in the supersonic diffuser with respect to the design shock wave impingement locations.

Bleed zone 1 was located just upstream of the shock-wave impingement on the cowl. Bleed zone 2 was located just upstream of the second shock-wave impingement on the centerbody. Bleed zones 3 and 4 were located in the throat region on the cowl and centerbody surfaces, respectively, and provided a variation in bleed mass flow as the terminal shock wave moved in the throat. The tests reported in references 1 and 2 indicated that these locations would provide satisfactory control of boundary-layer growth. All bleed zones were drilled with holes with a diameter to ca.pture radius ratio of 0.01 25. Bleed zones 1 and 2 were drilled to provide a uniform porosity of 41.5 percent. Bleed zones 3 and 4 had an overall porosity of 20.8 percent. The bleed hole pattern in each zone

4

could be altered by filling the holes with a plastic resin material. The method used to derive the final bleed pattern was described briefly in the introduction and is more fully described in the test procedure section.

Bypass System

The bypass system was located on the cowl surface just upstream of the engine face (fig. 2(b)). Sufficient area was provided so that at all test Mach numbers, all of the main duct flow could be diverted through the bypass. The largest bypass area was required at Mach number 2.0. The cowl surface was drilled with holes with a diameter to capture radius ratio of 0.01 25 to provide an overall uniform porosity of 41.5 percent (fig. 2(b)). The final bypass area included a correction for the effective area of the holes and was based upon the work reported in reference 7. The flow passed through the porous area into a p1enu.m chamber which was divided into three separate zones (fig. 2(b)), thereby reducing the possibility of recirculation of the flow. The flow then passed to the free stream through an exit which varied the flow rate from no flow to the maximum possible through the porous area.

Vortex Generators

Vortex generators were believed necessary to avoid high total-pressure distortion at the engine face. The generators, which were slightly taller than the anticipated height of the boundary layer, were located just downstream of the throat in order to induce the mixing action where the boundary layer was relatively thin. The spacing between them was chosen to ensure uniformly mixed flow at the engine face with a minimum number of generators. These considerations were based on previous experience (refs. 1 and 2). Other design details were based on the work reported in reference 8.

INSTRUMENTATION

Pressure instrumentation consisted of total- and static-pressure tubes as well as static-pressure orifices. The position of all moving parts was determined by the use of calibrated potentiometers. Six total-pressure rakes were provided at the simulated engine face. Each rake had six tubes spaced so as to provide an area weighted average total pressure. The tube spacing is shown in the sketch in table 2. Static-pressure rakes (fig. 2(a)) were located near the main duct exit and were arranged to give an area weighted average static pressure. Static-pressure orifices were located in single opposing rows along the top internal surfaces of the cowl and centerbody. They extended to the end of the subsonic diffuser. Boundary-layer rakes were located as shown in figure 2(b). Measurements from a seven-tube total-pressure rake, at the beginning of the throat, were used to evaluate the performance of the supersonic diffuser. Four-tube total-pressure rakes each with a single static-pressure tube were mounted in the centerbody bleed ducts. Four of these rakes were mounted in the outer duct and three in the inner duct (fig. 2(b)). Pressure orifices were located in the bleed and bypass plenum chambers. To evaluate additive drag for the transonic tests, four rakes were installed at the position of maximum centerbody diameter. Five total-pressure tubes and two static-pressure tubes were included on each rake. Both total- and static-pressure tubes were located to give area weighted average pressures.

5

TEST PROCEDURE

The investigation was conducted in the 8- by 7-foot, 9- by 7-f00t, and 11- by 1 1-foot test sections of the Ames Aeronautics Division Wind Tunnels. The transcnic Mach number range was investigated in 0.1 increments between Mach numbers 0.6 and 1.3 and the supersonic range was investigated in 0.25 increments between Mach numbers 1.5 and 3.5. Data were obtained at angles of attack of 0", 2 O , 5", and 8".

Supersonic Test

Attempts were made to redace the bleed mass flow in the supersonic diffuser (bleed zones 1 and 2) by closing rows of holes ill each zone. Only consecutive rows of holes were closed at either the forward o r aft end of the porous areas so that the overali porosity of the open area remained at 41.5 percent. Reducing the bleed through zones 1 and 2 in this manner caused the inlet flow to beconie unstable mr:king it difficult t o maintain a started condition at or near the design contraction ratio and, i n addition, resulted in relatively poor performance. In the throat region (bleed zones 3 and 41, distributed patterns of varying overall porosity were attempted by closing alternate rows of hoIes. The best performance was obtained by concentrating the throat bleed as far upstream as possible and by having three rows of holes open on each surface (fig. 2(b)). For the selected bleed pattern (fig. 2(b)), a contraction rztio that produced the best performance was deiemined experimentaliy. Three levels of bieed mass flow were then selected by varying only the throat bleed exit settings ( z o ~ ~ e s 3 and 4), because reducing the supersonic bleed exit settings was detrinlelltdl to the pcrform.ance. These correspond to bleed exit settings A, €3, and C oa the plotted data. Exit setting A represented the maximum bleed mass flow that could be removed for the sclccted bouildary-layer blecd-hole pattezn (holes choked). Exit settings B and C represented progressively reduced bleed mass-flow ratios. All reported data were obtained with these bleed exit settings and the selected bleed paltcxn.

Data wer-2 obtainzd a t sllyersnnic Mach numbers and at angle of attack with various fixed bypass exit settings. The exit settings were different for each Mach number and were selected to yield a range of bypass mass flows fro111 0 to the point where all the mass flow available at the engine faace could be diverted through the bypass. Most of the data at angle of attack and at off-design Mach numbers was recorded with the centerbody positioned to provide nearly the maximum coniraction ratio (lowest throat Mach number) a t which the inlet would remain started.

Transonic Pest

Transonic resu!:s (hQw = 0.6 to 1.3) were obtained with the bleed exits open and closed. The bypass exit was always closed for this Mach number range. The mass flow entering the inlet was deternlined frori, mt:asmements from the rakes mounted at the maximum centerbody diameter. Bleed and engine-face mass flows were not measured independently because of insufficient pressure ratio to choke the exits.

6

I

MEASUREMENT TECHNIQUES AND ACCURACY

The following table presents the estimated uncertainties of the primary parameters:

Parameter Accuracy +0.005 -1-8.005 20.02 +Q. 10" rt0.2 +o.oos k0.02, a = 0" and 2" 20.02, a = 0" and 2"

At angles of attack of 5" and 8", mass-flow ratios (mJm,) and (mi/m,> may be in error by +0.050 or more because of increasing flow distortion with increasing angle of attack. The uncertainties of all the parameters except mass-flow ratio are believed to be well established on the basis of experience gained through previous tests in the Ames Aeronautics Division Wind Tunnels.

Each system used for measuring the bypass and boundary-layer bleed mass-flow rates was calibrated as follows: Each exit was varied from fully closed to fully open; the flow rates calculated from the pressure measurements and geometric flow areas in each duct were compared with simultaneously measured incremental changes in main duct mass flow. Appropriate calibration factors for the effective flow areas for ezch bleed duct and for the bypass duct were thus determined. For bleed mass flow, this technique is believed to give the stated accuracy since, even though the absolute magnitude of the main du.ct flow is known only to 20.02, small changes in the main duct flow (0.10 or less) can be measured with much greater accuracy. Calibrations of the bleed flow ducts were made only at Mach number 3.5, but the results were believed to be equally valid at other Mach numbers as long as the exits were choked. Because the bypass involved large quantities of flow, calibrations were made at all Mach numbers, and i t is believed that bypass mass flow was about as accurate as the main duct mass flow. No calibrations were made at angle of' attack, but data are believed to be about as accurate a t 2" as at 0". At larger angles, increasing flow asymmetry could cause some circumferential flow in the plenum chambers which was not accounted for in the calibration procedure, and the accuracy would thus deteriorate. This procedure was used with apparent success in similar calibrations of the inlets described in references 1 and 2.

As stated previously, bleed and main duct mass flow were not measured in the transonic range because of insufficient pressure ratio to choke the exits. The inlet Inass flow in this speed range was determined from rake measurements at the station of maximum centerbody diameter (additive drag rakes) . These measurements were used to calculate the inlet mass-flow ratio and the total-momentum change from the free stream to the rake measurement station. They were also used with the pressure distribution on the centerbody, and a friction drag term, to compute the additive drag as explained in reference 9.

7

RESULTS AND DISCUSSlON

The results of the investigation are presented in figures 5 to 46 and in tables 2 and 3. The inlet theoretical mass flow at 0” angle of attack is shown in figure 5 for all of the supersonic Mach n.umbers at which data were obtained. The mass flow plotted as a function of the location of the cowl lip was obtained from a subroutine of the computer program described in reference 5. The results obtained at Mach number 3.5 are shown in figures 6 to 21. The results shown on each of these figures are discussed in the following sections. Figures 22 to 37 show similar data €or all other supersonic Mach numbers (3.25 to 1.55) but are not discussed because the discussion of the data obtained at Mach number 3.5 will suffice. Figures 38 to 43 summarize some of the results obtained throughout the supersonic Mach number rangc (3.50 to 1.55). -4 discussion of these results is included. Transonic performance is shown in figures 43 to 46 and these results are also discussed.

Bleed mass Row is used as a parameter for most of the plotted supersonic data instead of the more conventional mass flow at the engine face because i t is believed to be more accurate. At 0” angle of attack, the mass flow at the engine face is obtained by subtracting the bleed and bypass mass flows from the theoretical mass flow enlering the inlet. Oata a t angle of attack are shown as a function of mass flow a t the engine face.

Data in tables 2 and 3 are from the individual t u b a mounted at the engine face for the supersonic and trmsonic speed ranges, respectively. The sketch at the beginning of table 2 shows the location of each tube.

Table 4 is an index to the figures. Most of the off-design results include data for each Mach number tested. The quantities in parentheses were r,ot vaned for the data shown in each figure.

Perforrnance at M, = 3.50 Without Bypass

Supercriiicuk pcr.Ji’rr;wnce-- Maxinmn pressure rccovery does not necessarily occur at the nlaximuin inlet conlraciien ratio Lhat can be achieved without unstarting .the inlet. This is shown in figxes 6 and 7 . Supercritical performance for various positions of the cowl lip and bleed exit setting A is shown by figcre 6 , mcl the cowl lip positiol; is related io inlet contraction ratio by figure 7. Pressure recovery is virtually unaffected for the small range of cowl lip positions (x/R)lip = 2.830 - 2.848. Mcwever, at (x/R)lip = 2.835, the forwardmost position of the terminal shock wave was achieved without unstarting the inlet; hence, a higher pressure recovery and higher bleed mass-flow ratio resulted. For all cowl settings, total-pressure distortion remains low (about 8 percent or less) as long as the tern~inai shock wave moves within the confines of the porous bleed area in the throat (zones 3 and 4). Moving the termi!lal shock wave downstream of the throat ‘bleed region causes a rapid loss in total-pressure recovery, a rapid r ise in total-pressure distortion, and no further change in bleed mass-flow ratio. The forwardmost position of the cowl lip shown in figure 6 , (x/R)fiip = 2.82S, is nearly the maxirrwm cxmtraction ratio that could be obtained without unstarting the inlet and also represents a position for less than maximum performance.

In an effort to increase the efigine-face pressure recovery for a given amount of bleed, diffcrent amounts of bleed back pressure were investigated for the cowl lip position giving maximum pressure recovery, (x/R)lip = 2.835. Three combinations of bleed back pressure were

8

selected, as described previously, and the results are shown in figure 8. These data represent the range of performance available for the selected boundary-layer bleed hole pattern. Restricting the throat bleed reduced the supercritical operating range of the inlet; that is, the combination of the change in bleed plus the change in pressure recovery over the useful operating range was greatly reduced. The range of performance was 89.5-percent recwery with 21.5-percent bleed to 84.. 5-percen t recovery with 14-percent bleed.

Distortion- Total-pressure distortion was low (less than 8 percent) for all bleed exit settings over the useful operating range. The low level of distortion was attributable to the mixing action induced by the vortex generators and the fact that the final diffusion Mach ncmber was low (about 0.15). The radial distortion profiles are shown in figure 9 for each bleed ex.it setting. AU profiles are similar, but a progressively lower level of pressure recovery occurs as the bleed exits are restricted. In all cases the distortion from any single rake is about equal to the total distortion.

Bleed mass flow- The components of the total bleed of figure 8 are shown in figure 10. The bleed flow through each zone is plotted as a function of totabpressure recovery at the engine face. Most of the variation in supercritical bleed flow results from the change in the throat bleed (zones 3 and 41, although near maximum recovery changes also occur in the flow through bleed zone 2. The plenum-chamber pressure recoveries associated with these bleed flow are shown in figure 1 1 . These recoveries together with the bleed mass flow are necessary to assess the drag penalties associated with the boundary-layer removal system. The higher bleed pressure recoveries (zones 3 and 4) occur when the bleed flow rates are highest and ate caused by higher internai duct pressure. This is fortunate because higher pressure recoveries reduce the size of the bleed exit ducting required as well as provide a potential for lower bleed exit momentum drag coefficient.

Static-pressure distribution- Theoretical and experimental static-pressure distributions on the cowl and centerbody are shown in figure 12(a). T h i s figure shows the distributions for the most upstream location of the terminal shock wave (x/R = 4.10 - 4.20) fha.t can. be achieved without unstarling the inlet. Sabsonic flow occurs downstream. of the point where the static-pressure rise (p/p,) is about 40. Only partial agreement was obtained between the theoretical and experimental pressure distributions. The locations of the shock-wave impingement arc: in good agreement if' allowance is made for the fact that the experimental results were obtained with the cowl lip translated 0.025 x/R forward of the theoretical design position. Even so, the experimentally measured pressure rises across the impingement iocations are higher than predicted with the computer program. The combination of the difference in impingement location and the effect of the boundary-layer displacement thickness could partially explain the discrepancy. The measured pressure rise of 3.03 across the second impingement point on the centerbody includes the rise through the terminal shock wave. Figures 12(b) and 12(c) show the experimental pressure distributions as the terminal shock wave is withdrawn progressively downstream.

Now profiles- The effectiveness OF the boundary-layer removal system in controlling the boundary-layer growth is shown by figure 13. Pitot-pressure profiles upstream of, betweer;, and downstream OF each porous bleed area are shown in addition to a profile across the inlet throat. Near the inlet throat (x/R 4..20) boundary-layer height was tb.e same (h/R = 0.020) on both cowl and centerbody. On the cowl side of the flow passage, proEles upstream and downstream of the throat bleed (x/R = 4.1 8 and 4.38) show that the boundary layer is well controlled. However, on the centerbody side, near the same axial location, the tlicluiess of the low pitot-pressure region adjacent to the wall increased. This increase occurred over the relatively short distance x/R = 4.200

9

to 4.225 and was believed to be caused by a rapid turning of the flow 'in t'nis region (about 12"). This phenomenon could represent either an increase in boundary-layer thickness or a local reacceleration of the flow.

Attitude sensitivity with fixed geometry- The previous discussion has considered only the steady-state performance a t 0" angle of attack. Also of importance is the sensitivity of the inlet to sudden changes in the approaching flow conditions such as might be caused by gusts. A gust could suddenly change the local angle of attack by 2" to 3" before the inlet could respond with a change in geometry. Figure 14 shows the inlet tolerance t o changes in angle of attack. In this figure the total-pressure recovery a t 0" angle of attack has been plotted for three contraction ratios (positions of the cowl lip). The indicated angles ((Xuns) represent the limiting angle of attack, at various points along the supercritical operating cwves (starting from 0"), to which the model can be pitched, with no geometry change, without unstarting the inlet. As an example, with the inlet operating at (x/R)lip = 2.840 and maximum pressute recovery, the angle of attack can be changed only to 0.3" without unstarting the inkt. If the pressure recovery is reduced by moving the terminal shock wave downstream, the angle of attack to which the inlet can be pitched without unstarting increases to 2" and does not challge as the shock wave is moved farther clownstream. Decreasing the contraction ratio (increasing (X/R)lip) increased the tolerailce to changes in angle of attack, It should be noted that the data were acquired with a system capable only of slowly changing the angle of attack so that a sudden gust was not simulated.

Performance at M, = 3.50 With Bypass

Supercriticul performunce-- For steady state, on design operation, a matched inlet-engine cornbinatjon probably will require little or no bypass airflow. However, the bypass system can serve t o stabilize the position of the terminal shock wave when transient disturbances are encountered, assist in restarting the inlet in the event of an unstart, or permit establishment of stable inlet flow in the event of serious engi.ne malfunction. The effect of the bypass on the principle performance parameters as a function of mass flow at the engine face is shown in figure 15. Each curve shown corresponds to a fixed bypass exit opening. Small quantities of bypass mass flow have little effect on maximum pressure recovery while the larger quantities tend to reduce the maximum pressure recovery. The distortion remaixs low (about 5 percent) near the maximun pressure recovery for each bypass exit setting. Figure I 6 shows the same principle performance parameters as figure 15 plotted as a function of bleed mass-flow ratio. Subtracting the sum of the mass-flow ratios of figures 15 and 16 horn the mass-flow ratio of 1.000 entering the inlet gives the bypass mass-flow ratio.

Penalties-- 111 order to assess the drag penalties associated with the bypass system, the pressure recovery in the bypass plenum chamber must be known as well as the quantity of bypass mass flow. Figure 17 shows the bypass plenum-chamber pressure recovery as a function of bleed mass-flow ratia. Bleed mass-flow ratio is used as a parameter because the data can be plotted conveniently with zn expanded scale compared io that required if mass flow at the engine face were used. Comparison of the bypass plenum-chamber pressure (fig. 17) with the total-pressure recoveries at the engine face (fig. 1.6) shows that the total-pressure loss of the bypass air in flowing from the main duct through the perforated walls to the bypass plenum chamber increases with increasing bypass mass flow. This is an unfmorable effect because reduced pressure recovery tends to reduce the available momentum of the bypass air and increases bypass drag coefficient.

10

Effect on distortion- At or near maximum pressure recovery, the total distortion parameter docs not change as the bypass mass flow increases. However, the shape of the total-pressure profiles a t the engine face does change. Figure 18 shows the total-pressure recovery profiles at the engine face for two typical rakes on opposite sides of the duct (see figure in table 2). A profile is presented for each of the maximum pressure recovez?/ points shown in figure 15. Bypass mass-flow ratios of 0.02 or greater tend to improve the pressure recovery on the cowl side of the flow passage but at the expense of the pressure recovery on the centerbody side. The distortion is almost totally radial except for the case of maximum bypass where the distortion is mainly circumferential. The following paragraph gives a possible explanation for the increase in pressure recovery on the cowl side of the flow passage when the bypass is open.

Figure 19 shows the effect of bypass mass flow on the surface static-pressure distributions. Along the upstream portion of the bypass area (x/R = 5.025 to 5.175) on the cowl, the static pressure generally decreases as the bypass mass flow increases, while downstream of this location small amounts of bypass mass flow (up to about 6 percent) increase the surface static pressure. This indicates that flow separation may be occurring near x/R = 5.175 and is probably being eliminated by small amounts of bypass. On the centerbody no evidence of flow separation was iound so, as might be expected, increasing bypass decreased the static pressure. The apparent separation point on the cowl coincides with a relatively rapid change i n local surface slope and could be reduced or eliminated by recontouring the cowl surfac-c.

ArzgZe of uttack-- Figure 20 sllows the maxirnllrn pressure recovery and associated blccd mass-flow ratio and distortion as a function of bypass mass-flow raiio ;It angles G f attack of O", 2", 5O, and 8". These data are important because of the question of whether or not a system that performs well at small angles of attack will operate satisfactorily a t larger angles. Data at (3" angle of attack were obtained with a number of fixed bypass exit areas, and the data at 2", S o , and 8" were obtained with no bypass and with two other bypass exit areas. Flags indicate bypass exit openings maintained at a11 angles of attack. Increasing angle of attack decreased bypass mass fiow becalm internal duct pressure was reduccd. The largest bypass opening was sufi-icient t o diver t all the main duct flow through the bypass a t angles of attack of 0" and '2" (half-Tilled symbols). At any given angle of attack, changes in bypass mass flow have only small effects on the ~ n a x i ~ n u n ~ pressure recovery, bleed mass-flow ratio, and distortion. However, changes i n angle of' a t h c k r<!sillt i n large changes in these quantities. The cause of these changes will be discussed in the following paragraph.

Znkt-e~gine nurtching- Figure 21 shows the supercritical performance of the inlet at angle of attack with zero bypass and with a fixed bypass exit opening. Data with bypass are i1:cludeci because the normal mode of operation at 0" angle of attack may require a small amount of bypass. All data were obtained at contraction ratios slightly less than tkat which would cause irilet unstart. For reference, a constant corrected weight flow line has been added. Maintaining a started inlet at angle of attack requires a reduction in the inlet contraction ratio (cowl translation) which iilcreases spillage llow and distortion and decreases internal duct pressure and bleed mass flow. reduced contraction ratio and correspondingly higher terminai shock Mach number with a more nonimiforrn profile partially account for the reduced pressure recovery and increased distortion. These effects result in a change in engine face corrected weight flow. In the usual operation the bypass flow must be adjusted to satisfy the engine demand. At 2" angle of attack the engine requirements can be met without additional bypass. At 5" angle of attack, about IO-percent bypass is required, and at 8", about 30 percent. However, a t 5" the distortion may be unacceptably high.

I.. I"-

Off-Design Supersonic Performance

Data €or Mach numbers from 3.25 to 1.55 are presented in a form identical to that for Mach number 3.5. No analysis of these data (figs. 22 to 37) will be made since the trends and analysis presented for the data at Mach number 3.5 will, in general, be sufficient for understanding the off-design data. Table 4 indicates the type of data contained in these figures.

Maximum Pressure Recovery Performance at Supersonic Speeds

Figures 38 and 39 show maximum pressure recovery data obtained throughout the supersonic Mach number range. Included with the maximum pressure recoveries are the associated distortions and bleed mass flows without bypass at angles of attack up to 8", and with bypass at 0" angle of attack.

Figures 38(a) to (c) show the maximum performance throughout the supersonic Mach number range f x three bleed levels and no bypass. in general, at 0" angle of attack, maximum pressure recovery increases with decreasing Mach number. At Mach number 2.25 the pressure recovery is somewhat lower, possibly because of the change in bleed. This is particularly evident on the curves for bleed exit setting A where the bleed at Mach number 2.25 is considerably less than that at the next higher Mach nu.rnber. The curves for bleed exit settings B and C show smaller reductions in pressure recovery and correspondingly smaller reductions in bleed mass flow. The reason for the low pressure recovery at Mach number 2.0 is not fully understood, but may be due to the relative misalinement of the cowl and centerbody bleed surfaces. At Mach number 1.75 pressure recovery improves more than 5 percent and coincides with the Mach number where the inlet becomes self-starting (no cowl translation required to ingest the terminal shock system). The distortion. is below about 10 percent over the entire Mach number range.

For 2" angle of attack, similar trends were observed. For angles of attack of 5" and 8", substantial decreases in pressure recovery and increases in distortion occur. As might be expected, the losses in performance with 'increasing angle of attack are less severe at the lower Mach numbers.

Figures 39(a) to (c) show the maximurn performance as a function of bypass mass-flow ratio for three bleed levels. The half-filled symbols indicate that all the inlet flow has been diverted through the bypass. Only small decreases in pressure recovery occur over relatively large ranges of bypass mass flow. Distortion remains acceptably low over the entire range of bypass mass flow. Data that show an increase in pressure recovery with small amounts of bypass are probably the result of control of local flow separation on the cowl. In these cases, further increasing bypass mass flow reduces pressure recovery, which i s caused by losses on the centerbody side of the flow passage (see fig. 34).

Unstarted Inlet Performance

Figure 40 shows, at 0" angle of attack, the cowl lip position that caused the inlet to unstart and the position that allows the inlet t o restart (see fig. 7 for relationship between cowl translation and contraction ratio). A cowl translation distance of about 0.65 x/R is required to achieve inlet restart at Mach number 3.5, and this value decreases until, a t Mach number 1.75 and below, the

12

inlet is self-starting; that is, no cowl translation is required to restart. Figure 41 shows the ef€ect of unstarting the inlet on the principle performance parameters. The circles represent maximum values of pressure recovery and related values of distortion and bleed mass-flow ratio prior to unstart. The half-filled circles represent these quantities after the inlet unstarts. The pressure recovery and distortion during unstarted operation are subject to wide varia-lion because of the unsteadiness of the flow. The generally low level of pressure recovery and high level of distortion indicate the low quality of the flow at the engine face during unstarted operation of the inlet. Figures 42 and 43 irdicate the changes that occur in the individual bleed Cows and plenum-chamber pressure recoveries when the inlet unstarts. These quantities aid in the assessment of the attendant drag penalties.

Transonic Performance

The performance in the Mach number range up to 1.3 is treated separately because additive drag is a major portion of the presentation, and bleed and bypass mass-flow measurements were not made. Data were obtained with the bleed exits both open and closed but the bypass was closed for all testing.

Figure 44 shows the data obtained up to Mach number 1.3 with the bleed exits open. Achievement of high pressure recovery and low distortion requires some increase in additive drag over the minimum .value as well as some reduction in the inlet mass-flow capability. The results obtained with the bleed exits oper. and closed are compared in figure 45. The conditions far relatively large mass flows show little change in pressure recovery. fiwever, at reduced mass-flow conditions, pressure recovery decreases when the bleed exits are closed. The distortion may be unacceptably high, particularly at the lugher Mach numbers, for the conditions at high mass flow.

Data obtained at angles of attack up to 8" and Mach numbers up to 1.0 are shown in figure 46. 'The data shown are for the intermediate cowl lip position, (x/R)lip = 4.330 and with the bleed exits both open and closed. Because of flow asymmetry i t was believed that inlet mass-flow measurements at angle of attack would be uw"11izble and no attempt was made to make these measurements. Therefore, since the inlet gegmetry was the same at all angles of attack, the data are based on the assumptiorr that the capture mass flow was the same as ai Oo. With the exception of 8" angle of attack, the data obtained with the bleed exits open show a continuing increase in pressure recovery with a decrease in mass flow. With the bleed exits closed pressure recovery increases and then decreases as the mass flow is decreased. As was the case at O", the distortion at angle of attack may be unacceptably high for conditions of high mass flow.

CONCLUDING REMARKS

A 20-inch capture diameter model of a mixed-compression axisymmetric inlet system designed for a Mach number of 3.5 has been tested. Some of the main conclusions t o be drawn from the results are as follows.

The total-pressure recovery measured in the supersonic diffuser was somewhat below theoretical predictions partiaUy, at least, because the oblique shock waves were stronger than

13

predicted. In addition, in the vicinity of the throat (x/R = 4.200 to 4.225), a rapid increase in the thickness of the region of low pitot pressure occurred on the centerbody surface. This increase corresponded to an area of rapid change in centerbody surface slope (1 2" turning between x/R = 4.200 and 4.225). It is believed that a more uniform throat pitot-pressure profile could be obtained by reducing the rate of turning of the centerbody surface.

At the engine face with zero bypass mass flow, relatively low levels of total-pressure recovery were obtained on the cowl side of the flow passage. With small amounts of bypass mass flow the pressure recovery on the cowl side generally increased and that on the centerbody generally decreased. The increase in pressure recovery on the cowl side was believed t o be due to the effect of the bypass in reducing local flow separation on the cowl in the vicinity of the engine face. Flow separaticn was thought to be caused by an excessive rate of change in local surface slope in the region of the bypass. The overall effect of increasing bypass mass flow (at a fixed-bleed mass flow) was a general reduction in average total-pressure recovery with little or no change in the average total-pressure distortion.

Control of the boundary layer was accomplished with four porous bleed areas. The boundary layer in the supersonic diffuser was controlled by bleeding just upstream of two internal shock-wave impingement locations. Bleed in the throat region resulted in increased bleed mass flow and total-pressure recovery at the engine face as the terminal shock moved upstream in the throat region. Altering the throat bleed exits and hence the throat bleed back pressures varied the characteristic performance curves of bleed versus pressure recovery. Attempts to vary performance by increasing the back pressure of the bleeds in the supersonic diffuser resulted in the inability to achieye an inlet contraction ratio sufficient fcr high performance. At the design Mach number of 3.5 the total bleed flow rates appeared to be large, but compared with the trend of the total bleed flow rates of the inlets designed for lower Mach numbers, they may not be excessive. At off-design Mach numbers, the total bleed flow rates appeared to be excessive compared to those of the inlets designed for lower Mach numbers. A large portion of the off-design bleed flow was removed through the forward cowl bleed which, at the design Mach number, was just upstream of the shock-wave impingement. For off-design operation this bleed area moved well into the subsonic diffuser, where the pressures were relatively high; consequently, more b!eed flow was remcved from the cow! sur fxe than was necesszry for good performance. If the forward cowl bleed were relocated downstream of the shock-wave impingement location at the design condition, where the pressures are somewhat higher, the expanse of the porous area could perhaps be reduced. This would result in about the same amount of bleed at the design Mach number, but should reduce off-design bleed.

Over the useful supercritical range and at low angles of attack, vortex generators, just downstream of the throat on both cowl and centerbody surfaces, appeared to be effective in maintaining low total-pressure distortioil at the engine face over the entire Mach number range. However, because the Mach number of the engine face was low at the higher free-stream Mach numbers (about 0.1 5 at & = 3.5), low distortion may be easier to achieve than would be the case for inlets designed for higher engine face Mach numbers.

14

If, at 0" angle of attack, the inlet was operated supercritically and at slightly less than the contraction ratio for best performance, the inlet remained started with changes in angle of attack of up to 2". Such operation reduced the maximum pressure recovery about 1 to 3 percent without changing distortion.

Ames Research Center National Aeronautics and Space Administration

Moffett Field, Calif., 94035, July 14, 1970

15

REFERENCES

1. Sorensen , Norman E.; and Smeltzer , Donald B.: Investigation of a Large-Scale Mixed-Compression. Axisymmetric Inlet System Capable of High Performance at Mach Numbers 0.6 to 3.0. NASA TM X-1 507, 1968.

2. Smeltzer, Donald B.; and Sorensen, Norman E.: Investigation of a Nearly Isentropic Mixed-Compression Axisymmetric Inlet System at Mach Numbers 0.6 to 3.2. NASA TN D-4557, 1968.

3. Cubbison, Robert W.; Meleason, Edward T.; and Johnson, David F.: Effect of Porous Bleed in a High Performance hisymmetric, Mixed-Compression Inlet at Mach 2.5. NASA TM X-1 692, 1 968.

4. Sorensen, Norman E.; Smeltzer, Donald B.; and Cubbison, Robert W.: Study of a Family of Supersonic Inlet Systems. AIAA paper 68-580; 9. Aircraft, vol. 6, no. 3, May-June 1969, pp. 1 84-1 88.

5. Sorensen; Virginia L.: Computer Program for Calculating Flow Fields in Supersonic Inlets. NASA TN D-2897,1965.

6. Muehn, Donald M.: Experimental Investigation of the Pressure Rise Required for the Incipient Separa t ion of Turbulent Boundary Layers in Two-Dimensional Supersonic Flow. NASA MEMO 1-21-59A, 1959.

7. McLafferty, G.: Pressure Losses and Flow Coefficients of Slanted Perforations Discharging From Within a Simulated Supersonic Inlet. UAC Res. Dept. pi-0920-1, 1958.

8. Taylor, H. D.: Summary Report on Vortex Generators. UAC Res. Dept. R-05280-9, 1950.

9. Sibuikin, Mervin: Theoretical and Experimental Investigation of Additive Drag. NACA Rep. 1 187, 1954.

16

Tab le 1 .- INLET COORDINATES

I: ENTERBODY COWL

X r - - R R

0 0

s t r a i g h t t a p e r -

.288

.318 -356 * 395 .436 * 479

~ 1 I I t i I i L ! L

S t r a i g h t l i n e

.870

.830

.812

.849

- 794 777 .762 * 749 .745 742 .742 .746 - 773 .809 .819 .824

--

.825 E- Engine - face r a k e s S t r a i . g h t l i n e

17

Table 2.- SUPERSONIC ENGINE FACE PRESSURE RECOVERY DATA, pt-/pt 03

The following include total-presscre recoveries from the individual tubes mounted a t t h e engine face. Other performance parameters are also included. The sketch below shows the location of each tube.

Rake 1

4

Engine-face pressure tube location looking downstream

Table 2.- SUPERSONIC ENGINE-FACE PRESSURE RECOVERY DATA, pt2/pt, - Continued

B l e e d e x i t s e t t i n g B

" TUBE NO. II RAKE II TUBE NO. I

0.051 P,/P, = 64.0

n TUBE NO. ~ -7

10.877 10.87310.875 10.882 10.863 0.845

- 10.879 -~ 10.87010.87810.864 10.844 0.840

RAKE 1

i- u . I n 0 . 8 5 5 1 0 . 8 4 2

3 n0.83510.818

I 5 b . 8 5 1 [ 0 . 8 3 7

3.50

'IIJBE NO 3

0 . 8 3 4

0 . 8 3 0

0 . 8 3 1

4 1 s 0 . 8 2 5 10. E19

0.835 10.819

0.833 10.818 ..

Ty =. o.oo

6 0.805

0 .802

0 . no2 .

19

Table 2.- SUPERSONIC ENGINE-FACE FRESSURE: RECOVERY DATA, pt2/pt, - Continued

Bleed e x i t s e t t i n g B

RAICE NO.

2

4 6

"

II TUBE NO. I

I RA!! II TUBE NO. NO.

~ % ] ~ ~ & ~ ~ l 9 10.823 1 0.800 1 4 0 . 8 4 4 0 . 8 2 8 0 . 8 2 0 0 . 8 0 8 1 0 . 8 0 7 1 0 . 7 9 9 ~ 2

5 ~ 0 . 8 3 9 ~ 0 . 8 2 5 ~ 0 . 8 1 5 ~ 0 . 8 2 3 ~ 0 . 8 1 4 ~ 0 . 7 9 7 ~ 6

I II 1'IBE NO.

.. .

6 . .u

0 . 8 6 3 [

0 .8641

0.870n

- ..

NO.

2

4 6

"

."

"

3 ! .. ..4

0.86010.865

0.86210.869

0.85510.857

10.877 IO. 875

14.87810.883

l0 .85910.868

20

Table 2.- SUPERSONIC ENGINE-FACE FRESSURE RECOVERY DATA, pt2/pt, - Contil~.ued

Bleed ex i t setting B

1 1 0.818

3 10.794 1 .. 5 - 10.811

0.782 0.791 0.786

0.86610.873l0.845 -.841]-.84410.846 ..

0.851.10.86510.863 . . - . .

TUBE NO. E .? " ! .847 0

.815 0

.826 0

4 c). 828 0. 037 0.826

"_ .. ~. . 5 2

0.833

0.850 0.841

" " ._ . 3 0.824

0.837 0.834

~- - - "

.832

.818]

.836 . . I IO. 831 3.50 a = o.oo mo/mm = 1.000 M, =

TUBE NO. II I/ mF: NO. I "

0.81410.819[ 0.80510.82810.8368 2 ~~0.81410.81~+ 0.81710.8181 0.8l.410.812!0.808[ 10,-816[Q.817 0.81910.8221 0.82310.831/0.833/ 6 I! 10.81810.816

2 1

Table 2.- SUPERSONIC ENGINE-FACE FRESSURF: RECOVERY DATA, pt./pt, - Continued Bleed exi t s e t t i ng B

. . .

:. 6 ' 1 NO. . . . . _. . .

1.8401 2 10.845 .814n 4 10.806

. 8 3 8 ! " 6 . . - 10.846 . . . . . . . . .

1. 3. ! !...

10.846 10.832

10.83910.850 . . . .

1 5 [ 0.854

10.836 . . .

10.853

- 1 "3 lo. 847

10.810

10.857

I k .

10 856

10.827

IO. 869

I

' > - 6."J 0.824 1 o .836 1 . .

0 . 8 2 1 1 . . . . . . .

1

10.850 . . . L 5 1 0 . 8 0 0 [ 0 . 8 0 7 "" lO.82110.834 .. -

6 0 . 8 6 3

0.847

0.569

..

. "

1 5 84.6 1 .:

10.852

~ 10.842 .. I

1 5 . 10.858

IO. 835

10.872

I RAKF II TUBE NO. TUBE NO. 1 NO.

2

4 6

...

" "r- -' _. z q o . 8 2 4

. . . . . . 1 .__

0.80710.807 ...

0.809 0.81.1

.......

' t . - [ . - 5

. . . . . . I

... . . .

0.816 10.. 830

0 . 5 0 6 [ 0 . 8 0 9

0 .807 0 .806

~ . . " ~.

3.53 a = 5.0* ... M, = "" ~

II RAKE I I W E KC. I ...

!.

10.71 1

10.665

lo . 707

0.70210.715 " ""

I

0.68210.694 .~ . .

2 2

TaSle 2.- SUPSRSONIC ENGINE-FACE PRESSURE RECOVERY DATA, pta/pt, - Continued

B leed ex i t setting B

. . .. . . . ..

mm NO.

1

3 5

"

.. .

TUBE NO.

- 2 1 3 [ 4 0.71410.72110.716

0 .65410.67210.684

0 .64710.65510.669

n f,.?,,, 10.650

110.648

II TUBE NO. I RAKE NO.

2

4 6

.

. " .

"1 ~ _ _

I TUBE NO. I M: NO. "

2

4 6 ..

~~ . .

C . 3 . J 4 NO. u 1..

1 10.662

T "

2. 0.632

0 638

0 .633

.~ 4 0.645

0 I 636

0.642

~~

. . .. 0.705

0 .644

0 . 6 3 4

".

0.67510.690L0.688

0 . 6 3 6 ] 0 . 6 3 9 ~ [ 0 . 6 4 0

0 . 6 3 6 1 0 . 6 3 8 [ 0 . 6 3 4 "..

TUBE NO. PAKE / I TUBE NO. NO. 1 ]

4 0 . 4 3 4

6 1 0 . 4 5 :

2

0 .625

0 .436

0 . 4 4 3

- . -.

. .. -

0.557 1 0.489 1 0.490 1 ~~

I I 'PUBE NO. I I RAKE II TUBE NO.

I N O - .i. 1 . ! - ? - , I . - .,3 ! 4 . "

[ 1 n0.557]0 .61710.6501 0 .644

I 3 00.435[0.k401 0.4611 0.492

[ 5 10.4351 0.4441 0.4701 0.507

23

I

Table 2.- SUPERSONIC ENGINE-FACE FRESSURE RECOVERY DATA, pt./pt, - Continued

Bleed ex i t setting B

I II TIJBE NO. TUBE NO. I

0 . 4 1 2 . 0 . 4 1 2

0 .40910.410

RAKE: TUBE NO.

NO. ; 1 2 1 3 1 4

T[JEiE NO.

3- .L - 4":. 1.-5 I "6 ,

0,87110.883 ]0 .89510.366 1

3.25 M, = _____". a = o.oo mo/m, = 0.965 mbp/mm = 0

-. -

ptJpt, = 0.856 mbl/m, = G.151 = " 0.082 PJP, = 44.7 "" -

TUBE NO. . , . . ... _. - -.-

I NO. T I- . 4 .. -

1 lo. 878 10.886

Io. 890

I " 5 10.854

10.870

-~ "" 7 - I . !4

=i= G . 832 ---- ~- 0.86810.881 -

0 842rO. 851 10 .875

3

0 . 8 6 4

0 .869

0 .875

2

0 .323

2 . .

10.860 . .~

2

4 6 .. -

0.323

0 .879 0 .879 ~- 0,86810.88110.870 -

0 842rO. 851 10.875 . .

G . 826 G . 826

TIBE NO. I I Eu.Kfz II TVBE NO. II RAKE II 3

0.820

_ I ~ . --

2

0 , 8 1 9

0,-818

0.828

.-

_ _ 4 .- . "~

""

0.837

0 .833

0 .818

0 .835

0 .834

24

T a b l e 2.- SUPETSONIC ENGINE-FACE FRESSURE RECOVERY DATA, ptz/pt, - Conttnued Bleed e x i t se-tt ing B

M, = 3.25 a = o.oo mo/m, = 0.965 ?nbP/m, = 0.05

RAIm NO.

1

3 5

I.

. ... .

- 2 7 0.8311

0.8321 .. .

0.8361

TUBE NO. n 7 0.8351

0.8541

0.8431

""

0.875

0.845

m 0.885 0.9031 0 .844 0.86911

0.85710.885u

n u 1.

0 0.830

10.832

IO. 835

10.845

" 10.860

TUBE EO. 1 - 3 T

0.8561

0.8661

0.8831

M, = 3.25 a = o.oo 0 . 9 6 5

'rnE NO.

1 1 ]0.81610.81710.8211----

I 3~ 10.81610.823]0 .84310.854

I 5 10.81410.816[0.827[ . . ..... 0.837

""

. . .. . .. ".

0.83310.8571

@ . 8 6 6 [ 0 . 8 8 4 ]

0 . 8/15 I 0 .-87-1!

2

827

833

850 . . . ~

-w 0.869 0.88010.908

mbp/mw -- 0 . 0 5

PJP, = - 4 3 . 7

'TUBE NO. 1

3.25 a = 0 . oo mo/m, = 0.965 M, = ~ ~ ~ ~ _ _ _ _ " mbp/m, = - 0.05

0 . 0 7 2 P,/P, = 41.8

n TUEE NO. "1

TUBE NO.

0.8261 0.8321 ---- 0.8331 0,8501 0.865

0.82910.8411 0.838

5" 0.830 0.870

0.850

0.0341 0.880n

0.855 0

RAKE NO.

2

4 6

ll I I ~ l 3 1 p + I 5 6 ,

TGBE NO.

~~0.82610.835]0.853/0.858 10.866 0.880.

~ ~ 0 . 8 2 5 1 0 . 8 3 0 ~ 0 . 8 4 3 ] 0 . 8 5 5 . . ~ ~~ 10.871 0.885

10.52710.83410.84910.864 10.881 0.898

25

Table 2 " - SUPErRSOISIC EPU'GIPE-FACE FRESSUi RECOVERY DATA, pt,/p - Continued

B l e e d e x i t s e t t i n g B

to3

- - 0.20

= 43.6 "

1 . .. . . . . . . . " . . .

TUBE NO. ..

. . . ".

0.

0.

0.

. .

. " _ = . 3 0 842

0 .839

0.847 ...........

4 .. .. c . 354

0 .839

0 .863 ..... ........

....... " . . . . . ........

TKEE NO. 2

797

796

3C1 " .

1 3 1 ' i . .

10.81410.823 10 .830]0 .841

10.830 10.843 1 - 1 -

0.817

0.834 10.839 10.852 "" " " . .

3.25 Yo = _ _ _ _ " . a = 0.0' mo/mm -- 0" 962" mbp/mm 0 . 4 2

~.

3.25 a = o.oo -. mo/m, = 0 . 9 6 5 ...... ...... M, = ___,_ " .

TUBE NO. 1

26

I

Table 2.- SUPERSONIC ENGINE-FACE PRESSURE RECOVERY DATA, pt2/pt, - Continued

B l e e d e x i t s e t t i n g B

0.965 mbP/m, = 0.33

0.076 P2/Pw = 39 -8

RAKE n TUBE NO. N o m u " - 1 ! 2 1 3 . I 4 no.747~o.742~o.747~----

3 n0.73710.74910.758)0 .754

5 uO.73810.74510.75710.756

5 0.755

0.761 0.761

TTJBRF: NO. 2 1 3 1 4

0.80510.8061----

0.80810.81510.816 0.80710.809]0.805

5 0.800 0.820 0.810

TUBE NO. I

. . " ' n B E NO.

0.823 1 ---- 10 845 0.86010.863 10.838 0.86910.865 . . .. 10.838

0.852 I 0.818 1 0.817 1

'IIJBE NO .

. . .. TUBE NO. 1

2---!. 3

0.8121 0.822 0.797~l0.799 0.8151 0.840 . .

27

Table 2 .- SUPKRSOI\IIC EIVGINE-FACE FF?ZSSuRE RECOVERY DATA,

Bleed exit se t t i ng B

M, = 3.25 a = 5.0' m o b , = mbp/m, = - ""

- Pt2!Pt, 0.724 mbl/m, = 0.129 Apt2 = 0.147 P2/P, - - -___ 38.0

.... - . .. II I' "n'" ' . . . . . . . . . . . . . . . . . . . . . . . .

TUBE NO. '' ' '" I

~ ...

0.77910.7801

@.71310.735[

0 .?19[0 .738]

.... - ....... A

. . . . . . - .- , . . . . .n ..... - .............. I. ..._ ..... .. ......... TIJl3E NO.

"~ -

2

0 .491

0.473

0 a 489

... ".

....

TUBE NO. ~~ .

3 0.512

0 .503

0 .515

" . -

. ".

...

6 1 - . :.:. i o . 5 3 4

0 .521

10.530 . . . .

"

2

0,.461

(3.4 49 0 .461 ... I_

28

T a b l e 2 .- SUPERSONIC ENGINE-FACE FRESSURE RECOVERY DATA, pt2/ptw - Continued

B l e e d e x i t s e t t i ng B

M, = 3.00 a = o.oo mo/mw = 0.926 mbp/m, = 0

pi? I 3 I 5

~

1

0.881

0.891

0.885

TUBE NO. RAKE NO.

2

4 6

.. . .

TUBE NO.

2 7 3 4 5 1 6 -

.90110.921 0.900 0.883 0.855

.897[0.906 0.892 0.869 0.849

.91010.92510.899 0.873 0.851'

5 0.887

1

0.900 . ." - .

0.8631

-. 0.8501

0.8521 ~~

0

0

0 -

0.869

0.879 - 0.891

0.897

M, = 3.00 a = o.oo m o b w = 0.926 mbp/m, = 0

TUBE NO. TUBE NO.

2 1 3 1 4 1 5 6 - 0 .883~0 .90910 .883~0 .866 0 .835

0.885lO.895lO.874lO.848 0 .830~

0.89010.910]0.883 10.856 0 .831 '

1

0.873

0.876

0.867

~~ . 2 1 3 1 4 0.883]0.9031----

~-

5 0.862

0.847

0.857

1

0.883

0.875

0.876

0.8401 2

0.8311 4 0.8310 6

0.88310.89810.878

0.88010.90510.883

M, = 3.00 a = o.oo mo/m, = 0.926 mbp/m, = 0

Ft;Jpt, = 0.850 mbl/m, = 0.150 aptp = 0.104 P,/P, = 29.7

RAKE II W E NO. IRAKE TUBE NO.

1.87210.869 10.882 10.854 B.833 10.807

1.872 10.868 10.873 10.848 10.823 0.803 ~

1.871 10.873 10.882 10.857 10.825 0.805

NO. 1 1 - 1 2 I 3 1 00.86710.86610.892

3 10.870]0.863]0.871

5 10.86710.86710.889

4 "" 10.83510.812 I 2

0 .851 10.817 10.803 I 4 0.856 10.82810.805 I 6

""

0 .851

0.856

3.00 a = o.oo mo/m, = 0.926 M, = Qp/rnw = 0.04

TUBE NO. TUBE NO. 1 II II T-" I

2 4 3 0.885 ---- 0.901

0.877

0.8681 0.8771 0.886

0.906 0.895

3

0.899

0 .901

0.904 "

5 0.916

0.908

0.904

29

Table 2 .- SUPERSONIC ENGINE-FACE FRESSURE RECOVERY DATA, pt./pt, - Continued

B l e e d e x i t s e t t i n g B

PJP, = 31.6 ..

TUBE NO. I 3

0 .873

0 .885

0 .888

c " .

"

4 0.885

0 .898

0 .902

. "

.. .

" . .. .

1 . 5 . IO. 888

10.885

I O . 888

.. - I

.~ -1 0.8811

0 .8851

0.8851 -

--- . E 6 1

0.83;1

0.8343

0 . 8 3 4 )

r==- I RAKF: II TUBE NO.

__

G

0.860

0 . 8 6 1

0 .861

.~ - - . . . . - -. .

30

I

T a b l e 2.- SUPERSONIC ENGINE-FACE FRESSURE RECOVERY DATA, pt2/pt, - Continued

B leed ex i t s e t t i n g B

%a= 3.00 01 = o.oo m,/m, = 0.926 m-bp/m, = 0.45

- pt2/ptw 0 . 8 6 4 mb&, = 0.169 npt2 = 0.085 PJP, = 32.2

RAKE NO.

1

3 5

. *.

I / RAKE It TUBE NO. 1

0.88910.895u 4 ~ 0 . 8 4 1 ~ 0 . 8 4 4 ) 0 . 8 5 3 ~ 0 . 8 7 1 ~ 0 . 8 9 0 ~ 0 . 9 0 1 ~-

0.866_10.8860 6 ~ 0 . 8 4 0 ~ 0 . 8 4 2 [ 0 . 8 5 3 ~ 0 . 8 7 3 ~ 0 . 8 9 6 ~ 0 . 9 0 9

M, = 3.00 a = o.oo mo/mw = 0 . 9 2 6 mbp/m, = 0- 44 ~- - - Pt2/Ptw = -0.854 m t , l / P ~ = 0 . 1 6 2 Apt2 = 0.095 P,/P, = 31.9

N o - . u 1 - 1 10.830

3 10.829

5 10.827

2

0 .838

0 .833

0.830

'I'UBE NO.

3 I 4 I 5" 0.8541---- 10.887

0 .84410.85110.877

0 .83510.83410.847

0 . 9 0 4 1 2

0.8841 4 0.8681 6

TUBE NO. 1 0 . 8 3 1

0.830

0.827

I ~ ~ l ~ 1 2 1 3

' r n l

1" 1 nO. i82 [0 .79310.820

I 3 b . 7 8 1 1 0 . 7 9 1 [ 0 . 8 1 3

I ~ 0 . 7 8 0 ~ 0 . 7 8 6 ~ 0 . 8 0 1

NO. .. a I 5 1 6 8

---- io . 856 10.854 I 0 . 8 3 1 [ 0 . 8 4 3 [ 0 . 8 4 6 1

0 . 8 1 3 ] 0 . 5 4 1 1 0 . , 8 4 4 ~

TUBE NO. 1 NO.

2

4 6

31

Table 2.- SUPERSONIC ENGINE-FACE FRESSURF: RECOVERY DATA, pt2/p - Continued

Bleed ex i t s e t t i n g B

t,

M, = 3.00 (r = 2.0° mo/m, = "" mbp/mcu = O

mbl /m, = 0.181 Apt2 = 0.096

TUBE NO. . .

- Pt2 /P tm = 0.881

0.842 0.8551 0

= 31.8

I I TUBE

. "_ 3 1 0.9241

0.9131 - .

0.8881

"

3. I 4 -1 5 ;

.8971---- 1;. 898

.87810.902 IO. 902

.86O] 0.8781 0.893

1

0.876

0.846

0.868

. . -~

0.8961 0.8661

0.8991 0.8701

0.897]0.860]

TUBE NO. I MF: II TUBE NO. 4

"_ .815

.825

1 5 . 10.892

10.843

1; . 893,

! 6 . 10.899

10.903

10.865

= 30.8 ___"

I RAKE II TUBE N O . I T [ - - * I i

798 10.806 1 .800 10.810 1 ,804 Io. 814 ]

1. 0 79810 - 795 b . 7 9 8 1 0 . 7 9 8 " .. . ~ .

3-00 a = 5.0' M, =

2

0.737 -.

-~ 0.707

0.735

" .. ". . . 6

0 .751

0.727

0.749

- . ... - .

0 .7540 .77210 .781 -

32

Table 2 .- SUPERSONIC ENGINE-FACE FRESSURE RECOVERY DATA, pt2/pt, - Continued Bleed ex i t s e t t i n g B

RAKE NO.

1

3 5

"

1 TIJRE NO. TUBE NO. 1 6 NO.

0.735 0.753[0.7761---- 10.80710.8171 2 O . 6 8 6 ~ 0 . 6 8 6 ~ 0 . 6 9 8 ~ 0 . 7 0 6 ~ 0 . 7 1 4 ] 0 . 7 2 6 ~ 4

0.75310.7761---- 10.80710.817~ 2 0 . 6 8 6 1 0 . 6 9 8 1 0 . 7 0 6 ] 0 . 7 1 4 ~ 0 . 7 2 6 ~ 4 0 . 6 8 6 1 0 . 6 9 4 ] 0 . 7 0 3 1 0 . 7 1 8 ~ 0 . 7 2 6 ~ 6

0.735 0.686

0.6821 0 . 6 8 2 ~ 0 . 6 8 6 ~ 0 . 6 9 4 ~ 0 . 7 0 3 ~ 0 . 7 1 8 ~ 0 . 7 2 6 ~ 6

M, = 3.00 a = 5.0' mo/mm = "" mbp/m, = 0- 56

- pt2/pt, = 0.669 mbl/m, = 0 .124 npt2 = 0.218 P,/P, = 27.0

,,!y I I - I

TUEJF: NO.

3 1 4 0.7371 ---- 0.65010.647

0.64810.643

TUBE NO.

2 1 3 1 4 1 5 6

0.64510.64910.66210.677 0 .684- 0 .647[0 .650[0 .658 0 .663 0 .674

0.64810.65410.670 0.687 0.691'

1 2

0.710

0 .651

0.646

5 0 .770

0 .649

0.645

0.645

0.650

0 .642

0.7881 2

0.6528 4 0.6481 6

0 a = 8.0'

[El 1 1 2

1.- n 0 . 6 6 5 1 0 . 7 2 1 I 1 0 . 4 9 0 1 0 . 4 9 8 I n 0 . 4 9 3 1 0 . 5 1 0

TUBE NO. II = n TUBE NO. N o - ! 1 I 2 1 - 3

o . 5 5 3 10.523- n0.493 10.515 10.555~0.,571

6 ; 4 1 5 2 00.545 10.52910.555 0.546 0.587 0.584

6 10.551 10.530 10.557 10.601 0 .591 10.536

3.00 a = 8.0' m o b , =

= 0.584 %l/m, = 0.137 P t 2 = - 0.375 P,/P, = 23.3

"" mbP/m, = 0.04 M, =

Ft,/Ptm

IRAKEll TUBE NO. 1

RAKE NO.

2

4 6

TUBE NO. I

I N 0 - U 1 1 2 1 3 ! 4

I i0.5231 0.5331 0.55d 0 .569

I 1 [0.65810.708] 0.7391 ----

[ 5 u0.5201 0.5311 0.544 0.557

1

0 .565

10.524

I O . 568

0.5861 0.59011

0.5791 0.5930

33

Table 2 .- SUPERSONIC ENGINE-FACE FBFSSuRE RECOVERY DATA, pt,/pt, - Continued

Bleed exit setting B

" L ."

10.492

10.477

10.495

TUBE NO. "

RAKE TUBE NO. No. 1-

0.475 0.475

0.474 0.476

3.. 0 .721

0.478

0 .481 ~~

" NO. I 5 - 1 . 6 . .!" - .

~ 0 . 4 9 7 ~ 0 . 5 0 0 ~ 4 "

]0.50510.515n 6

M, = 2.75 .~ ~ ~ a = o.oo malm, = 0.852 mbp/% = 0

= 0.902 mbl/m, = 0.190 a p t 2 = 0.085 PJP, = 2 2 - 3

TUBE NO. 1 TUBE NO. B ~~ 1

3 ! 1 5 . i . 6 . 1 0 .912]0.937]0.925]0.8831

0 .92110 .94510 .930[0 .894]

0 . 9 2 7 ] 0 . 9 4 3 ] 0 . 9 3 1 ] 0 . 8 8 4 ]

. -

2 10.878

4 10.871

6 10.874

0.889

0.889

0.899 IO . 869

% = 2.75 a = o.oo mol% = 0.852 -~ .

P,/P, = 21.6 ~

~ . ..

TUBE NO. " .

1 0.095

W E NO.

3 i ' no. 84510.

10 :85O 10.

[0-.-848 10.

0.8631"" ." " ~.~

0.88410.910

0 .87810.911

TUBE NO. I .~ L I " 5 1 P".!

L ~ . ~ ~ ~ L . . . ~ ~ ~ I IO.. 902 10.836 1

]0 .90610.826]

2

0 .828 .~

3 "

4

0 .&8 8

0,9 15

0.9.13

0.847.LO.885

0.842 . -. - . - 10 . . - 8 8 0

Table 2.- SUPERSONIC ENGINE-FACE FRESSURE RECOVERY DATA, pt,/p t, - C o n t i n u e d

Bleed ex i t se t t ing B

~ -

TUBE NO. TUBE NO.

[ 1 ]0.83910.84310.8521---- 10.91410.923[ 2 nO.84410.84710.857 0.88410.918 0.91;

I 3 ~ 0 . 8 3 9 ~ 0 . 8 4 0 [ 0 . 8 5 9 ~ 0 . 8 8 5 ~ 0 . 9 0 8 ~ 0 . 9 2 1 ] 4 10.84510.85310.873 I 5 ~ 0 . 8 4 3 ~ 0 . 8 4 8 ~ 0 . 8 6 4 ~ 0 . 8 8 4 ] 0 . 9 1 2 ~ 0 . 9 1 1 ~ 6 nO.84510.85110.864 0.89610.923 0.929

0 .919 0.90210.935

. ~ " -

I N O - U 1 1 2 I 3 ! 4 I 1 ]0.826,]0.82710.839]----

[ 3 n 0 . 8 3 0 [ 0 . 8 4 2 [ 0 . 8 5 4 [ 0 . 8 7 2

I 5 i 0 . 8 3 4 1 0 . 8 4 1 1 0 . 8 5 6 [ 0 . 8 7 4

TUBE NO. 2 - 1 3 1 4 1 5 1 6 -

.82810.83810.867 10.90810.918

-836 0 .85610.888 10.925 (0 .920,

.840 0.85310.879 )0.91410.929

TUBE NO. 1

10.86910.8981 2 n0.800 10.806

10.88510.9061 4 n0.806l0.815

I 5 b.805[0.81810.83510.847 10.88410.9101 6 b . 8 0 6 10.817

TUBE NO.

3 I I - 5 6 0 . 8 2 1 10.846 10.882 0.905 0 .844 10 .8754.909 10.908

10.834 10.860 10.902 10.914

2.75 a = o.oo mo/m, = 0.852 M, = %p/mw = P J 3

?t, /Ptm = 0.877 mbl/m, = 0.185 a - P t 2 __ -

0 , P L . PJP, = ~ 2 2 . 2

TUBE NO.

35

Table 2.- SUPERSONIC ENGINE-FACE FRESSURE RECOVERY DATA, pt2/pt, - Continued

B l e e d e x i t s e t t i n g B

RAKE NO.

2

4 6

." " - __ k . -. . - . - n0.825l0.829

n0.82510.832 10.83110.838

TUBE NO. . . .~

1 F N E TUBE NO. n r5 107867

10.896 Io.900

6 0.900

0.911 0.912

... . . .

~

NO. 1 2 3 " !,.. 5 " - 3 0.821 0.821 0.829 ---- 10.861

3 0.826 0.832 0.85010.86210.885

5 0.827 0.833 0.84210.85310.883

0.83810.848

0.84710.870 10.84810.864

TITBE NO.

3 1 4 1 5 0.80810.82910.856

2

0.793

0.803

0.793

M, = 2.75 a = 0.0'

&t$pt, = 0.858 q,l/rn, = 0.178 ~

RAKE TITBE NO.

= 0 . 3 1

= 21.7

1 6 1

3.8611

3.9191

3.916 1

RAKE NO.

n TUBE NO. L NO. 1

P 3 1 . 2 1 0.850 0.839 0.834

3 0.84710.8681 0.835

A 5 0.83510.8381 '0.832

4 1 5 6 0.905

0.914

0.866

~

5 .846

.877

.892

1

I O . 834 10.833 10.841 10.853 10

10.841[0.851]0.86110.878 10 ---- 10.890

0.873 10.902

0.840 10.852

2

4 6

2.75 a = o.oo mo/m, = 0.852 QP/m, = 0.30 M, =

et2/ptm= 0.838 %l/m, = 0.164 A P t 2 = 0.135 PJP, = 21.2

W E NO. 1 ')JIE[T NO.

2 10.815

4 6 00.816

___ 6

0.873

0.867

0.881

~~

.

~ ~.

RAKE TUBE NO. NO.

~~~ "" -

1 3 1 4 - 1 . ~ 5 - 2

1

0.82210.8271 0.83010.851 0.818 5 0.8341 0.85510.869 .~

0.818 '0 .816 3 0.8201 ---- ~. 10.846 0.814 0.813

- 3 I 4 ! 5 1 ..I 0.818[0.826l0.839lO.871]

0.82310.837 10.861[0.8941

0.83510.845 j0.865i0.9011 1

- 2

0.814

0.814

0.825

36

Table 2.- SUPERSONIC ENGINE-FACE F'RESSURE RECOVERY DATA, pt2/pt, - Continued

Bleed ex i t s e t t i ng B

M, = 2.75 a = o.oo mo/mw = 0.852 q P / m , = 0.26

- Pt2/Pt, = 0.801 mbl/m, = 0.123 npt2 = 0.151 PJP, = 20 - 1

TUBE NO. I TUBE NO. [ I

oo.760 10.763

00.761

.. 5 0.829 0.831 0.813

m n "0- 1 1

2 10.827

4 00.827 6 10.821

TUBE NO. TUBE NO.

- 2 1 3 1 4 5 , ~~

6

0.825]0.823]0.820 0.816 0.801 0.8261 0.83210.838k844 0.853~

0.825]0.82710.83410.844 0.850'

- 2

0.826

0.825 0.829

3 1 4 0.8291----

NO. 11 2 ..

[ 3 10.820 1 5 .n0.825

5 0.829

0.831 0.834

6 0.840

0.838 0.840

0.8291 0.827 0.83510.830

a = 2.0°

Pt2 /p tW = 0.858 q , l / m w = 0.160 APtp = 0.143 P,/P, = 20.8

I II I n 1 TlTBE NO.

3 1 4 0.8881 ---- 0.83710.859 0.85210.867

II TUBE NO. 1 NO.

2

4 6

1 1 -

lo. 810 10.821

IO. 829 _____

2

0.855 0.814 0.824

5 0.916 0.877 0.880

6 0.858 0.859 0.853 3.856 10.893 10.926 10.905 10.826 1

" a = 2.0°

2 3

0.827

0.817 0.835

0.843

0.822 0.857

37

Table 2 .- SUPERSONIC ENGINE-FACE FRESSURF, RECOVERY DATA, pt./pt, - Continued Bleed e x i t setting 3

M, = 2.75 a = 5.0' . . . - . .. mo/m, = "" . mbp/m, = O

""

TUBE NO. . -- . . "

0.69510.694j0.705 - . . . . . . .

0.670 10.67810.689

0.716 10.73910.759

0 .694 10.708-10.724 0 . 6 ~ 0 . 6 9 1 ~ 0 . 7 0 8 ~ 0 . 7 2 2 ~ 0 . 7 4 0 ~ . 7 5 9 I 1 I

" "_ " . ~ -" ~. . . . . -~ -

"

6

L642

. ~ - ."

0.630

0 .649 _- -

38

I

Table 2.- SUPERSONIC ENGINE-FACE FRESSURE RECOVZRY DATA, pt,/ptm - Continiled Bleed ex i t se t t ing B

TUBE NO.

3 1 4 0 . 9 0 1 I ---- 0.90710.934

0 .92410.949

5 0.952

0 . 9 3 1

0 .932

TUBE NO. 1

4 ~0.87710.89810.92910.95010.943 0.887 6 n0.879[0.90610.94010.95510.942 0.877

M, = 2.50 a = 0. oo mo/m, = 0 . 7 5 1 mbp/mm = 0

- p t 2 / p t , = 0.884 %l/m, = 0.164 npt2 = 0.129 P,/P, = 1 4 - 5

2

0 . 8 5 1

0 .863

0.877

TUEiE NO.

3 1 4 1 5 .~

O.8821---- 10.936

0.88910.92110.916

0 .91810.94110.891

RAKE NO.

2

4 6

._

U l b 1 nO.84010.8591

10.84910.8761

10.84310.8691

-1 TUBE NO.

0.89370.932 0 .915 0 .856

0.92110.94810.912 0 .849

0 .915 i 0 .939 io-

TUBE NO. 1 1. 2 1 3 I 4 1 ~ 5 - 1 6

TUBE No. 1 1 2 . ! 3 1 q 5 82910.86lTO.8861---- 10.885 0.827 I 2 10.829 10.879 10.91710.9080.851 10.797

82110 .85510 .879 ]0 .884 10 .856 82510.86810.91410.908 10 .84510.7931 6 10.818 10.85110.892 10.900 10.860 10.795

0 .8031 4 10.830 10.865 10.907 10.907 10.863 10.808

2.50 0, = o.oo mo/m, = 0 . 7 5 1 M, = %p/% = .om05

TUBE NO.

2 1 3 1 4 - 0.8511 0.8651 ---- 0.8591 0.8781 0.901

0.8701 0.8971 0.923 -

3.168 a = 0.118 P,/P, = . L 5 L

" B E NO. 1 P t 2 - -

5 0.930

0 .928

0.942 uO.85310.86810.90110.935 ~~ 0.94110.928

~0 .84610 .86210 .890(0 .926 0.93910.94C'

39

Table 2 .- SUPERSONIC ENGINE-FACE FRESSW RECOVERY DATA, pt2/ptm - Continued

Bleed exit setting B

&, = 2.50 a = o.oo mo/m, = 0.751 " QP/m, = 0.05

TUBE NO. R A K E n TUBE NO.

3. . 1 o. 8441 0.8551

0 ., 8641 _ _ ~

4 "

""

0.866

0.890 . ..

2 .".-.

.829

.839

.825

0.8341

F831(

P,/P, = 14.1

TUBE NO. ' . 1 1

TUBE NO. NO. 2 1 3 1 4

. I ~ . ~ ~ ~ - ~ . ~ ~ ~ ~ ~ . ~ ~ ~ ~ - - - - -

3 10.807. 0.82.5 0.858T0.873 ~~

5 10.816 0.857 0.90310.923 , .

"

5 0.869

0.870

0.906

213[ 4 ! 5 . I ..6 -1 0.83910.878~0.9~0[0.896~Cj.877]

0.8.53[0.885/0.89610.873[0.847]

0 . 8 2 7 ) 0 . 8 6 9 ] 0 . 8 9 4 ] 0 . 8 8 8 / 0 . 8 8 6 ] . . .

mbp/mm = 0.11

PJP, = 1.5.3

TUBE NO. - ~ - -- -

I r=- " - TUBE NO. , . - "

f- [O. 859

10.861 10 .-86 6

..

- 3 1 8731"

8841 0.

.. 87910. . ._

h "

."

89 9

902

"

2

0.864

0.869

0.879

"

2 1 0.86410. 0.870 10.

0.86310. .~

"

M, = 2.50

~~ "

a = o.oo .. ." ~ m h , = 0.751

I RAKE I1 T U B E NO. TUBE NO. I " -

2 "

3 0.859 0.881 0.865

\ 4 - ! 5 .! 6 1 10.888 ]0.91910.921] 10.916 p.943 10.933 I 10.896 1".9271".9401

"

2 1

0.85410

0.85610

0.840 0.854

0.846

A8L52L:"-- lo, .873]0.89110.

- .877]-0.908[0.

40

Table 2 .- SUPERSONIC ENGINE-FACE FRESSURE RECOVERY DATA, ptn/pt, - Continued

Bleed e x i t s e t t i n g B

RAKE NO.

1

3 5

., . .

TUBE NO. II RAKE II 1 I 2 ~I 3.. I 4 I,, 5 I 6 . -. II . No. ~ .. u11 ..,

0.81210.81710.8331---- IO.89O[O.923O U0.812I

0.822~0.842~0.858~0.875]0.907~0.916~ 4 10.8201

0.819~0.837~0.856~0.881~0.925~0.929~ 6 no.8131

TUBE NO.

~~ ~ 2 , - 1 3

0 . 8 1 8 [ 0 . 8 3 4 0.863 0 . 9 0 7 1 0 . 9 2 6 ~

4 1 5 I 6 ~

0 .93110 .930 0 .83210 .858 l0 .893 0.82210.84410.87310~921~0.937A

RAKE I NO. 1 3 I 5.

TUBE NO.

3 1 4 0.8601 ----

0.8661 0.877

0.8601 0.874

5 0.88Z

0.895

0.895

2

0 . 8 5 2

0 .847

0 .869

TUBE NO.

M, = 2.50 a = o.oo m o b , = 0.751 rnbp/m, = 0 . 1 8 ..

1 5 ---- 10.861

0.862 10.883

0.857 10.891

~

6

0 . 9 0 2

0 .907

1 0 . 9 2 6

NO.

2

4 6

IF I 1

1 3 I 5

1

3.789

3.791

3.792

TUBE NO.

!'. . ,3. I 0.8131 0.8471 0.864

0.8141 0.8511 0.881

5 0.858

0 . 8 6 8

0 . 9 1 6

6

0 . 8 6 6

0 .860

0 .916

rn NO.

2

4 6

II 1

0.794

0 . 7 9 8

0 . 7 9 2

Table 2.- SUPERSONIC ENGINE-FACE FRESSURE: RECOVERY 3ATA, pta/pt, - Continued

B l e e d e x i t s e t t i n g B

". . . . .

TUBE NO. .. - a

3 0.852

0.855

0 868

..

" 1 -

0.8551 ---- 10.862

0.85910.865]0.885

.~ 0.8521 0.85210.865

- pt2/pi;= 0.848 mbl/m, =

0.167 npt2 = 0.136 ..

-

TUBE NO.

4 0.8?01G.830~0.842

6 H 0.83010.837]0.842 - ..

5

I). 858 0.851

2.50 cr:= o.oo mo/m, = 0.751 ~ "" . mbP/m, = 0.33 M, = -

TUBE NO. I TUBE NO

li

824 IO. 847

0.814 10.831 13.856

0.806 . . 10.816 10 .!37

n c ~~

.792

.807 "

0 798 ~

42

Table 2.- SUPERSONIC ENGINE-FACE FRESSURE RECOVERY DATA, pt2/pt, - Continued

B l e e d e x i t s e t t i n g B

M, = 2.50 01 = 2.0' mo/m, = "" mbp/m, = O

- pt2/pt, 0.870 mbl/m, = 0.178 Apt2 = 0.137 P2/P, = 14.4

TUBE NO. n NO. i/ .. 1 I ! 3.. ! 4 ! 5 I " 5 4 1 [0.85510.86410.8821---- 10.92710.8940

3 ~ 0 . 8 3 5 ~ 0 . 8 4 0 1 0 . 8 5 2 1 0 . 8 6 4 1 0 . 8 7 2 ~ 0 . 8 5 8 ~

5 ~ 0 . 8 3 6 1 0 . 8 4 3 1 0 . 8 5 4 1 0 . 8 6 7 ) 0 . 8 7 7 1 0 . 8 ~

-n TUBE NO. No- 11 1 I 2"-1 3 I 4 I 5 1 . , 6 - 2 ~ 0 . 8 5 0 ~ 0 , 8 8 7 ~ 0 . 9 1 5 ~ 0 . 9 2 7 ~ 0 . 9 1 7 ~ 0 . 8 4 7 ~

4 ~ 0 . 8 2 7 ~ 0 . 8 2 8 ~ 0 . 8 4 4 ~ 0 . 8 6 2 ~ 0 . 8 7 9 ~ 0 . 8 6 1

6 ~ 0 . 8 5 7 ~ 0 . 8 8 9 ~ 0 . 9 1 7 ~ 0 . 9 4 3 ) 0 . 9 1 6 ~ 0 . 8 2 4

2

0 .852

0.809

0.810

TUBE NO.

3 1 4 0.8701 ----

0.8171 0.820

0.8211 0.825

1

0.829

0.810

0.826

TiTBE NO. I 2 1 3

0.84410.865

0.81310.825

0.84210.863

415 0.894l0.927

, 0.841 10.863 10.892 [0.927

~~

~~~~

0 .888

0.933

% = 2.50 a = 2.0° mo/m, = "" q p / m , = 0 3 4 ~

I ) t4pt t , =0.807 mbl/m, = 0.178 . - 0.178 P,/P, = 13.6

I N 0 - u . 1 I I 3 I I 5 ! - " d N o . 1- 1 n0.81710.829]0.844 ---- 10.86110.8731 2

1 3 nO.79310.80010.802 [ 5 ~ 0 . 7 9 8 [ 0 . 7 9 8 [ 0 . 8 0 3 1 0 . 7 9 5 1 0 . 7 8 7 ~ 0 . 7 7 8 1 6

0 .79310 .787[0 .7781 4

!I 1 1 1 2

no. 802 l o . 800

10.795 10.794

nO.79110.800

TTJBE NO. I

0.799 10.810 10.818 10.831 I 0.800

0 .804

TUBE NO. . . - I

2

0.760 0.720

0.776

TUBE NO. I . . 3~-

0.773 0.733

0 798

0.7g01.78510.732/ 0 .737 0 .728 0 .709

0.818 0 .788 0 .724

43

111 I I I I II I I IIIIIII111111111 l11111111.111111 111 .,...,I- .,...,, I,

Table 2 .- SUPERSONIC ENGINE-FACE FRESSURE RECOVERY DATA, pt,/p - Continued t,

Bleed exit setting B

% = 2.50 a = 5.0' mo/m, = "" q P / m , = 0.09

PJP, =

. ! - NO.

0.76310.

0 .77510 .

0 .75410.

0.178

n TUBE

VL! ~ 0 . 7 3 1 ] 0 . 7 3 4 1 0 . 7 4 3 ]

n 0 . 7 9 2 [ 0 . 7 6 3 [ 0 . 7 5 7 [

10.78910.76010.7601

mbp m~ / = 0.44

p /p, = 13.9 2 Pt,/Pt, = 0 .708 q , l / m , = 0.138 npt2 = 0 . 2 6 1

RAKE NO.

2

4 6

RAKE II TUBE NO. TUBE NO. 2 I 3 I I s . ! ?

0 .683]0 .687]0 .69410 .707[0 .706

0 .68210.68410.68410.683~0.681

3 . 6 8 7 [ 0 . 6 9 3 [ 0 . 6 9 7 [ 0 . 7 0 8 [ 0 . 7 1 4

p 0.683

4 1 5 ---- l o . 856

. -

3 0.827

0.687

0.685

. . ' 2 i

0.793

6 0 .853

0 .671

0.678

1

0 .683

0.685

0.686

0.685 0.68110.676

0 . 6 7 9 ] 0 . 6 8 1 0.684 -

TUBE NO.

k * d & O [0.659(;646

0.795 0 .842 0 .836 ---- 10.78110.669

1 - - I

10.654 10.637

10.682 10.652 1 5 u0.668j0.667]0.67310.667 10.66410.648 "

2.50 a = 8.0' M, = mo/m, = "" QP/m, = 0 .07

RAKE TUBE NO. NO. 1 2 I 3 4

" .. .

' 1 0 . 7 8 5 0 . 8 4 0 . 8 4 4 ----

3 0.659-0.657[0.661[ 0.659

0.81910.789ll 2 0.68210.66710.66710.669 0.740

0.652

0.741

44

I

Table 2.- SUPERSONIC ENGINE-FACE FRESSURF: RECOVERY DATA, pt2/pt, - Continued Bleed exit setting B

% = 2.50 a = 8.0° mo/m, = "" mbp/m, = O-37

- Pt2/Pt, = 0.662 mbl /m, =O. 1 2 7 a p t 2 = 0.361 P2/P, = 13.2

RAKE NO.

1

3 5

"_ ~

TUBE NO. TUBE NO. 1

- pt2/pt, = 0.896 mbl/m, = 0.155 a p t 2 = 0.095 PJP, = lo.o

TUBE NO.

.1~! ! 3 ! 4 ! 5 1 6 0 . 8 6 7 ~ 0 . 8 7 2 ~ 0 . 9 0 1 ! 0 . 9 2 4 ~ 0 . 9 3 3 ~ 0 . 9 0 7

NO. I . - 4 2

2 n0.86710.881

4 10.864lO.886

6 10.87010.891 ~_____

TUBE

- 3 I 0.9111

0.9211

0.902 I NO.

4 0.915

0.920

0.943

M, = 7 ~ . 7 5 a = 0 mO/m, 0.628 mbp/mm = 0

iS t aP t , = 0.882 q,l/m, = 0.149 ~p~~ = 0.087 P,/P, = 9 . 8

= n TUBE NO. m n TUBE NO.

-No. 1 1 . 1 2 1 3 4 1 5 - 1 6 6 No- 1 1 1 2 1 3 I 4 - ~ 1 5 1 ~ 0 . 8 4 8 ~ 0 . 8 6 2 [ 0 . 8 8 3 ~ 0 . 9 0 8 ~ 0 . 9 1 9 ~ 0 ~ 8 9 0 ~ 2 ~ 1 0 . 8 7 5 [ 0 . 8 8 5 [ 0 . 8 9 8 [ 0 . 8 9 3

0.847 0.898 5 ~0 .864 [0 .878 ]0 .88810 .88810 .88510 .850~ 6 ~ 0 . 8 6 6 ~ 0 . 8 9 0 ~ 0 . 9 0 0 ~ 0 . 9 1 1

0.850 0.909 3 ~ 0 . 8 5 9 ~ 0 . 8 7 8 ] 0 . 8 9 8 ~ 0 . 8 9 4 ~ 0 . 8 8 4 ~ 0 . 8 5 1 ~ 4 10.861[0.88810.91210.924

0.858

2.25 a = o.oo mo/m, = 0 .628 mbp/m, = 0 M, =

Ft2/ptm= 0.850 mbl/m, = 0 . 1 3 5 P t 2 = - - 0.122 P,/P, = 9.4

RAKE II TUBE NO. TUBE NO. NO.

1

3 5

L

1 ! * ! 3 1 4 1 5 1 2 1 3 1 4 1 5 0.81010.830] 0.864 0.88710.88810.8521 2 10.81410.83810.870

0.82010.8461 0.884 0.90310.86910.799n 6 nO.81210.83710.86710.88910.876

0.876 0.896 0.8021 0.8221 0.854 0.86810.8641 0.8111 410.81410.84110.872 0.875 0.892

-~ -

161 0.817

45

Table 2 .- SUPERSONIC ENGINE-FACE FRESSW RECOVERY DATA, pt,/pt - Continued

Bleed e x i t s e t t i n g B

0 3

n . .

TUBE NO.

M, = 2.25 ~

a = 0 . 0 ~ - mo/m, = 0.628

TITBE NO.

0.8711 0.8931 0.916

0.8731 0.8941 0.912

ll II TIJBE NO .....

1

0.080- PJP, =

TUBE NO. I

2.25 a = o.oo .. . m d m , = 0.628 M, == -__ - ~ -

= 10.2 "_ -

46

Bleed e x i t s e t t i n g B

M, = 2.25 a = o.oo m,/m, = 0 .628 q P / m , = 0 . 1 4

II TUBE NO. II RAKE NO.

2

4 6

II TUBE NO. I

' I D E NO.

" ! 3 1 4. 0.83310.86510.893

0 . 8 2 4 ] 0 . 8 3 6 [ 0 . 8 4 4

0.82110.83210.836

TUBE NO. I

M, = 2.25 . a = 0 .oo mob, = 0 . 6 2 8 Q,/III~ = 0 .26 .____

W E NO.

! 2... I 3 ! -.4 ! 0 . 8 4 3 ! 0 . 8 4 8 ! 0 . 8 6 1

10.85410.86610.876

10.847 10.858lO. 858

10 .880[0 .8841

10.86810.8751

T'UJ3E NO.

4 10.346 10.850 10.864 10.880

6 10.851 10.857 10.861 10.875

I RAKE II TUBE NO. II RAKE I I ~~

4 1 0.. 867 I o .87810.888[

.862/ o. 878[1

. 8 8 6 ] 0 . 8 9 5 [

NO. I1

2

2 [10.82810.839

4 ]10.83310.840

6 1 0 . 8 3 2 ~ 0 . 8 3 c

TUBE NO.

47

Table 2.- SUPERSONIC ENGINE-FACE PRESSURE RECOVERY DATA, pt2/pt, - Continued Bleed exit setting B

M, = ~- 2.25 a = o.oo m o b , = 0.628

NO.

2

4 6

" " -

- .. .

. .- r n . 1 " . . . - . [ 0 . 8 0 1 ] 0 , 8 0 7 [ 0

[ 0 . 7 9 8 1 0 . 8 0 4 ] 0

[ 0 . 8 0 2 [ 0 . 8 0 8 ] 0

. . .I - TUBE NO.

~- - - 3 - .. -. . I .? - .82710.841

.82110.850 -

.816[ 0 ,832 . ~ " i . ; ~ ~ .

. ..

M, = 2.25 " a = o.oo mo/m, = 0.628 . .~ rubp/% = 0.47

pt2/pt, =0.853 q,l/m, = 0.158 apt, = 0.040 P,/P, = 9.8 -

RAKE TUBE NO. NO. 1 2

.~ ! 3. - 1 4 ".

1 ' 0 . 8 4 7 0 . 8 5 0 1 0 . 8 5 ~ [ 0 , 8 5 0

. 3 0.853 0.85310.85510.851 5 m0.848.0.849[0.85010.845 ~~ - .,_ .

5 0.851

0.846

0.847 - ..

6

0.856

0.841

0.839 I . .

R A K E - n - - -

112 852 l o . 851

I

0 . 8 3 5 ~ 0 . 8 4 2 ~ 0 . 8 5 9 ~ 8 7 0 ~ 0 . 8 6 5 ~ 0 . 8 3 9

0.840 0 .848~0 .860~0 .862_10 .869~0 .842 . . . "

- _ l _ _ ."~ - .. " - .. ~ ~ ~

2.25 a = 2.0° - Do/% = "" -. . - QP/m, = 0 . 1 3

= 0 . 8 6 3 %l/m, = 0.147 A = 0. i 1 3 P,/P, = 10.1.

M, =

~ t , / p t m P t 2

48

Table 2 .- SUPERSONIC ENGINE-FACE FRESSURF: RECOVERY DATA, pt2/pt, - Continued

B leed e x i t se t t ing B

M, = 2.25 a = 5.0' m o h w = "" q P / m w = 0

0.816 mbl /m, =0.138 Apt2 = 0.158 PJP, = 9 - 0

TUBE"N0. 4 I TUBE NO.

e 3 . 1 4 1 . 5 1 2 3 1 4 5 6 0.8781 0 .893 [0 .88810 .870 [0 .804~ 2 10.86310.846 0.83910.8; '0.8370.785:

0.78510.79010.79110.79O10.773n 4 ~-10.779 0.779 0.780 0.785 0.772 0.764

0.78410.78310.785]0 .78610.769~ 6 10.858 0 .857 0 .864 0 .870 0 .845 0 .780-

TUBE NO. 8-n TUBE NO.

2 . r 3 / l ~ ! 5 ! 6 . N o ' n 1 ! 2 ! 3 ! 4 5 6

0.87210.888]0.89110.889~0.876~ 2 ~ 0 . 8 3 6 1 0 . 8 0 7 1 0 . 7 9 5 ~ 0 . 7 9 9 0.804 0.812-

0.76410.76710.77210.77610.777[ 4 n0.769]0 .77210.776 0 .780 0 .785 0 .789~

0.767 0 .770 0 .767 0 .770 0 .771 0 .840 0 .817 0 .806 0 .812 0 .819 0 .827 -

.. . TUBE NO

NO.

1

3 5

.I- .865 1 0 . .71610.

.714 10. . .

0.697 10.685

0.696 10.691

6 0.704

0.676

0.679

2 1 0 .75410

0 .693 10

10.761 10

2.25 a = 8.0° mo/m, = "" Inbp/% = 0.10 M, =

Pt2/pt,= 0 . 7 3 3 ' ~ 0 . 1 1 7 A P t * = " 0 260 P,/P, = 9-0

I T ? 1 - I 3 I 5

.-.I

F ["l 2 1 0 . 8 0 1

4 10.689

6 10 .801

2-1 3 0.73110.712

0.6851 0.682

0.74410.729 ~~ ~~

49

Table 2.- SUPERSONIC ENGINE-FACE FRESSURF: RECOVERY DATA, pt2/pt - Continued

B l e e d e x i t s e t t i n g B

rn

- . . . . . . 1

I 5 10 r8 6-6 10.883

-

10 I 866

; . . 6 i .-

- 0.8483

0.8581

0.8531

4 2 ', -.

0.

0.

0. ~-

927

YO8

9 24

I 1 !0.90710.9221 - 1 1

0.877

0.885 . .-

0.901 .. . .

..- , _.. ,

TUEE NO. .-

3 1 4 o.87n10.857

0.873 0.851

0.867 I. 0.852 . .

NO. 4 5.

0.849

0.838

6 0.816 1 0.8191

0.8191 .. .

0.853

0 860

0.852 . .

0.842 ., . .

mbp/mm = 0 . M, = 2.00 - . Q: = o.oo mo/m, = 0.518 .~=

" - . " . -.

RAKI;: I1 TUBE NO.

I. lo. 841. IO. 838 I" 0.830 ._ .

- " .

2 . .

0.879

0.865 0.87910.88210.8740.86210.833 . . . _" . . "~ "_ ~~- .. . " - "

1 RAKE I I TUBE NO " . "

6 0.859

0.870

0.855

-. ."=

"~ -.

3 0.899

0.915

0.902

." " " .

..

. " .

" . . . . .-

- .li .. ~,~ . Is . ~. . . ~

0.88410.868 .

0.89510.877 -. ~ ~

0.874 10.860 "" . .

" 0.87610.8661 _ _ 4 [L0.899[0.910 .~. "

0.8741 .~ 0.8671 "" "_ 6 10.907k.920 ". "-

50

Table 2.- SUPZRSONIC ENGINE-FACE FRESSURF: RECOVERY DATA, pt2/pt, - Continued

B l e e d e x i t s e t t i n g B

0 .518 mbp/m, = 0.05-

0.067 PJP, = 6.6 ~

1 TUBE NO.

No- . ~ - - II 1 ! 2 1 4 i . 5 . I . - 1 2 6 1 10.8971 0.8991 0.8791 0.86010.8531 0.8511 2 3 [ 0 . 8 8 5 ~ 0 . 9 0 1 ~ 0 . 8 9 7 ~ 0 . 8 7 4 ~ 0 . 8 6 0 ~ 0 . 8 5 4 ~ 4

i 5 ~ 0 . 9 0 1 ~ 0 . 9 0 6 ~ 0 . 8 8 9 ~ 0 . 8 6 2 ~ 0 . 8 5 5 ] 0 . 8 5 1 ~ 6 0 . 9 0 4 ~ 0 . 8 8 1 ~ 0 . 8 6 1 ~ 0 . 8 5 3 ~ 0 . 8 5 0 ~ I 1

2

0.873

0.872

0 .883

3 I . 4 0.87610.357

0.88710.872

0.88410.859

5 0.850

0.857

0 .851

'RJBE NO.

- 3 1 4 1 5 0.86710.875 10.891

0.84510.850 10.862

0.85710.862 10.874

a = 0 . oo

NO.

2

4 6

" i i 1 p

no. 850 l o . 863

10.835 f0.642

10.835 10.837

6 0.903

0.877

0 .899

. . .

..

10.851

10.850

0 .875

0.097 P,/P, = "h.5

ll TUBE NO. 1 TUBE NO. . . "

. .. " m p K E

..

5 NO.

0.8581 0.86211 2

0.8381 0.8511 4 0.9391 0.8461 6

1 1

1 3 1 5

n 1 ! - 2 I -3. .!-

0.804[0 .81610.82110

o . 8 0 7 ~ 0 . 8 1 7 [ 0 . 8 3 7 [ 0 - 0.8121 0.8271 0 . 8 3 8 b

3 .!. 0.8441 0.846

1 0 .814 0 .824 0.824 0 .827

51

Table 2.- SUPEBSONIC ENGINE-FACE FRESSURE RECOVERY DATA, pt2/pt - Continued

Bleed ex i t s e t t i n g B

03

TUBE NO.

.. -2- 1 .. ..!+-

0.79810.829

0.80810.837

0 ,78410.812

- - 1 1 5 I b - 1 10.83610.8381

] 0 . 8 5 4 [ 0 . 8 6 9 1

[ 0 . 8 3 3 ] 0 . 8 5 2 ]

I RAKE II TIJBE NO. - 3 ~ .~" .. 1 . 4 0.8181 0.830

0.7761 0.785

0.-7871 0.793

M, = 2.00 a = o.oo mo/mm = 0 . 5 1 8 mbp/m, = 0.27

j55,/ptm = 0.849 rq,l/m, =0.130 Apt2 = 0.075 - P,/P, = 6.8

I RAKE II TUBE NO. TUBE NO.

7 3 ! 4 1 5 - 1 6 . 0 . 8 3 3 ~ 0 . 8 3 3 1 0 . 8 4 8 ] 0 . 8 6 0 ~ 0 . 8 6 4 1 1 1 1

1

0 . 8 3 2

0 . 8 3 1

0.840 ~ ~~

3 , 0 . 8 2 9 1 0 . 8 3 1 1 0 . 8 3 9 ] 0 . 8 3 9 ~ 0 . 8 4 6

5 0 . 8 2 5 ~ 0 . 8 3 1 ] 0 . 8 3 8 1 0 . 8 4 3 [ 0 . 8 5 5

0.83310.83910.848 10.85810.889

0 . 8 5 0 ] 0 . 8 6 6 ] 0 . 8 7 2 1 0 . 8 7 5 1 0 . 8 7 5

TUBE NO. ~~

3 1 4 jzzJEi 0.82810.830

0.83110.837

5 !. 0.83710.

0,836 [ 0 .

0 .85210.

1-1 3 " . I 4 . - A- 1 -5 ~ .

[0.81710.82510.829 10.841 10.844

h 1 9 1 0 . 8 2 3 10.832 10.840 10.846

10.820 10.825 10.834 10.841 10.847

~ " " .

P -7 3 . 8 4 3 1

3.851 I 3.853 1

L -~ 0.824

0 .820

0.825

L

-

II TUBE NO. II TUBE NO. 1

0.807

0.80E

0.814

- ~- -

" ".

".

*..80810.814 0.824

0.8141 0.8281 0 . 8 3 4 0 . 8 4 5

0 .804 0 .8121 0182{0 .826

52

Table 2 .- SUPERSONIC ENGINE-FACE FRESSURE RECOVERY DATA, pt2/pt, - Continued

Bleed exit set t ing B

M, = 2.00 = 2.0° mo/mw = "" mbp/mw =

- Ptz/Pt, =O. 875 %l/m, =O. 1 3 1 Apt2 = 0 . 1 0 8 P,/P, = 6.6

" " c NO. 1 1 10.863

3 10.854

5 10.860

TUBE NO.

0 . 8 8 3 ~ 0 . 9 0 7 ~ 0 . 9 2 3 ~ 0 . 9 1 7 ] 0 . 8 6 0

0.86010.871]0 .869 10 .86410.841

0 . 8 5 9 [ 0 . 8 5 9 [ 0 . 8 5 1 ] 0 . 8 4 6 ] 0 . 8 3 1

W U TUBE NO. N o . u l 1 2 3 4 6 5

2 10.86910.895

0 .831 ' 0 .897 6 10.874 [0 .907[0 .914 10 .921

0 . 8 3 8 ' 0 . 8 7 4 4 [ 0 .851 ]0 .86710 .879 10 .891

0.83; 0 .898 0 .925 0 . 9 1 4 . - "

M, = 2.00 a = 2.0° mob, = "" mbp/mw = 0 . 1 8

Pt,/Ptw = 0 - 8 6 0 q l / m w = 0.130 Apt? = 0.104 P,/P, = 6 .9

RAKE TUBE NO. TUBE NO. ~~

3 4 5 6 2 1 NO. 3 6 5 4

I n 0 . 8 3 6 1 0 . 8 5 1 0 . 8 6 g l D , 8 8 3 ~ o . g o l ~ o . g 1 6 ~ o.919: o.910 o .884 o . 8 4 9 ~ 0 . 8 6 0 o.840

I 3 UO.83010.833]0.836]0.833 ] 0 . 8 3 4 ] 0 . 8 3 1 1 4 0.904 0 .907 0 .901 0.867 10.884 0.849 1 5 ~ 0 . 8 3 1 ] 0 . 8 3 7 ~ 0 . 8 3 6 ] 0 . 8 2 9 ~ 0 . 8 4 0 ~ 0 . 8 3 6 ~ 6~

0.869 0 .867 0 .853 0.836 10.854 . '0 .837

M, = 2.00 a = 2. oo mob, = Q P / m , = 0.38

P t ,&tm

""

= 0.828 %l/m, = 0.126 = - 0.107 PJP, = h.8

RAKE NO.

n TUBE NO.

ii t . ! * ! . . 3 ! m 10.8221 0.8391 0.8471 0.85710.8691 00.81710.822] 0.8201 0.81110.8131

[0.81510.8131 0.8171 0.81010.809I

6

0.881 0.792

0.792

TLTBE NO. 1 A .825]0.% .821 ] 0 .821

. 8 2 3 ] 0 . 8 4 4

5 3

, I , I .................. .""

Bleed ex i t setting B

" - - TUBE NO. n ... .- . . ~ . -._ I1

... ...

" .. I I I

0.821~0.816~0.80510.785~ " . . . _ _

" 0 . 8 0 9 1 0 . 8 0 3 l 0 . 7 9 ~ 0 . 7 8 0 ~ "_ . -. - " - ." . .

.~ I

3 .. 1 .4..? 1 . 2 - 1. I. -! 0 . 7 9 1 j 0 . 7 9 5 ~ 0 . 7 9 0 [ 0 . 7 8 3 ~

0 . 8 8 3 1 0 . 8 ~ 1 1 0 . 8 5 4 1 0 . 7 7 9 ]

0.89110.887 " . l0.855l0.786 . ". I

TUBE NO.

0 . 7 7 9 j 0 . 7 8 0 [ 0 . 7 7 8 ~ 0 : 7 7 7 [ 0 ~ 7 7 5 I .

- r - . - r - -

m, = "" mbp/m, = 0.35 . "

0.153 " P2/Pw = 6.7 - . . ...... "

TUBE NO. 1

54

......

Table 2 .- SUPERSONIC ENGINE-FACE FRESSURE RECOVERY DATA, pta/pt, - Continued

Bleed exit s e t t i n g B

0.106 Apt- = 0.262. P 2 h , = 6.4 -

TUBE NO.

. .. " " 1 TlTBR NO. TUBE NO. I ."

-. 3 . .

0.718

0.690

4 0.707

0.686

~" ~~ 3 ! 4 !.. 5

0 .87910.893/0 .894 . . ..

2

0 .753

0 .691

0 .752

0.9831 0.979

0 . 9 6 6 1 0.946

TlTBE NO. 1 3

0.932

0.935

0.933

__-

~~~

~~

0.896

55

. .. . . "

Table 2 .- SUPERSONIC ENGINE-FACE FRESSURE RECOVERY DATA, pt2/pt, - Continued

Bleed exi t s e t t i n g B

% = 1.75 a = o.oo mo/m, = 0.451 %PI% = 0 - pt2/pt, = 0.906 mbl/m, = 0.114 Apt* = 0.128 p2/P, = 4.6

M, = 1.75 a = o.oo mo/m, = 0.451 mbp/% = 0.03

- pt,/Pt, = 0.960 q l / m , = 0.139 npt2 = 0.074 P2/P, = 5.0

~ ~ $ p ~ , =0.938 mbl/m, = 0 .133 np t2 = 0.070 P2/P, = 4.9

56

Tab le 2 .- SUPERSONIC ENGINE-FACE FRESSURE RECOVERY DATA, pt2/p - Continued

Bleed ex i t s e t t i n g B

t,

M, = 1.75 a = 0 . oo mo/m, = 0 . 4 5 1 mbp/m, = 0 0 9

- Pt2/Pt, = 0.932 % l / m ~ = 0.132 Apt* = 0.062 P2/P, = 5.0

% = 1.75 a = 0 . oo mo/m, = 0.451 - " . ." mbp/mw 0.08

Ft2 /P tm = 0.909 %l/m, = 0.125 Apt? = 0.062 P2/Pw = 4.8 ." .. ~

T'ITSE NO. TUBE NO. 6

1.75 a = o.oo mo/mcu = -~ 0.451 mbP/m, = 0.18 M, =

T TUBE NO. I1 RAKE II 2

0 .946

0 .934

0.942

. ..

- ~-

TUBE NO. I

57

Table 2 * - SUPEXSONIC ENGINE-FACE FRESSURF: RECOVERY DATA, pt2/pt, - Continued

B l e e d e x i t s e t t i n g B

..... PJP, = 5 . . . . . .

TUBE NO. 1 ... - .... NO. n

.....

.^_ 3 ~.. I., i... 1 . _ ? . 1"16."i 0.947 10.958 10.957 10.91Ol

0 . 9 3 2 ~ 0 . 9 3 6 ~ 0 . 9 4 0 ~ 0 . 9 ~ ~ 5 ]

........ 0.93110.938'[0.937 - . - 10.939 1

... ,". .........

2 . 10.929 " 4 . . . . 10.926 .. i 6 ". 10.928 ~

F -~ . . . . ".- .. .- ............... ." n

..... " .. " ... - ........ - ... " TUBE NO.

58

" ......

Table 2 .- SUPERSONIC ENGINE-FACE FRESSURE RECOmRY DATA, ptn/ptrn - Continued

Bleed ex i t set-ting B

1 - n TUBE NO. TUBE NO. I [No- u 1 ! 2 1 3.. !, - 4 1

I 3 ~ ~ 0 . 9 1 4 ~ 0 . 9 2 2 ~ 0 . 9 2 9 ~ 0 . 9 3 7 ~ 0

I 1 n0.95610.96910.97710.976]0

I 5 [ 0 . 9 1 2 1 0 . 9 1 6 1 0 . 9 2 5 1 0 . 9 2 7 ~ 0

7 0.9681

0.9481

0.9401

~~

5 0.975

0.933

0.950

5 .982

.945

.937

4.

0.982

0.928

0.958

3 0 963

0.922

0.956 0.942

TUBE NO .. " " . i i 1 ! 2 1 3 ! 4 j 5 [ - ~ ~ N O . i j l I ~ .

1 nO.957iO.96610.97710.969 10.968[0.948[ 2 10.926 10.908

3 n0.870 10.869 10.86810.370 10.875 10.882 I 4 10,854 10.850

5 nO.85610.85810.86610.864 .. . .- \0 .870[0.870 I -. 6 . nO.911 lO.900 _____

.. .

TUBE NO. . . .

TUBE NO. 1 3

0.889

0 800

"

0.895 "4- 0.855 0.766 * [ 3 1 . 4 2

0.905

0.794

0.867 0.87310.867 ~~ 10.835 10.7841

59

_. ... .- ."

Tab le 2.- SUPERSONIC ENGINE-FACE FRESSURE RECOVERY DATA, pt2/p - Continued t,

Bleed ex i t s e t t i ng B

% = 1.75 a = 8.0' mob, = "" . q P / m , = 0.07

PJP, = 4.6 Apt* = 0 . 1 7 1 . .

NO. TUBE NO.

0 . 8 8 8 J 0 . 9 0 0

0.795 10.792

~~ 0.78610.792

M, = 1 . 5 5 a = o.oo mob, = 0 . 4 2 1 mbp/m, = o - Pt2/Pt, = 0.980 %l/m, = 0.132 Apt2 = 0.076 PJP, = 3 . 9

TUBE NO. RAKE TUBE NO.

0.962 0 .972 0 .988

! 5 .;I ...6 IO. 987 10.926

10.987 10.962

10.989 10.962

- 1 5 ..

7 0 . 9 9 5

10.989

10.994

~

2 4 I 3. I " ." 6 0 .968

0 .959

0 . 9 6 1

10.981

10.975

10.981

0.980

0 .993

0.995

1.000

1.000

1.000

0.995 4 10.970

6 10.970 0.996

~ ~ z / p ~ , = 0.973 %l/m, =0.127 Apt2 = 0.089 ~ P,/P;' 3 7

.- ~ -- " .. .~ I n 1

0.970

0.962

W E NO.

2

9.9 83

0 .985

0 .975 ~~ ~~

0.99110.982 -

TUBE NO. I RAKE TUBE NO. RAKE NO. 7 2 - 1 3 ." 4 5 6 No* 1

~ -~ ~- 1 0.960 0.97_010.970 0.958 0.974 0.930 2 - 0 . 9 7 3

3 . 0 . 9 6 6 0 . 9 . 7 d 0 - : 9 A 0 . 9 7 3 0 . 9 7 1 0 . 9 2 1 4 10.963

I 5 - 0 . 9 5 6 0 . 9 6 5 ~ 0 ~ ~ 9 7 2 ~ 0 ~ 9 ~ 1 0 . 9 7 4 0 . 9 2 2 , 6 110.963

3

0 .962

0.973

0 . 9 7 1

__- ~ ..

-~

2

0.977

0 .966

0.967 0.976 10.972 10 60

Table 2.- SUPERSONIC ENGINE-FACE FRESSURE RECOVERY DATA, pt2/pt, - Continued

B l e e d e x i t s e t t i n g B

M E n TUBE NO.

1 !0.95110.96510.972[----

. 3 ~0.95710.971~0.976]0.971

I 5 10.950[0.953j0.963[0.962

I1 n 1 10.962 10.955 I O . 961

TUBE NO. 1

I0:if%$2?Zj 10.963 0.971 0.981 0.992

j0.976 io.99o I

n ] 1 ! 2 ! 3 ! 4 [ 5 p N O * [ 1 1 [ 2 ] 3 ! 4 1 5 6 0 . 0 956 I 0.966 0.967

TUBE NO. = n TUBE NO.

---- 0.971 0.973 2 0.970 0.972 0.962 0.975 0.974 0.948

~0.96210.97210.97310.966 10.97610.975 4 10.946 10.955 10.965 0.969 0.970 0.972 ~

n0.943]0.95410.964[0.960 10.97010.9721 6 00.957 10.963 10.965 0.967 0.969 0.970

TUBE NO.

2 1 3 1 4 1 5

TUBE NO. 1 I

6 1 Nom /I 1 1 2 1 3

0.9571 2 10.97610.98810.983

0.9421 [0.95710.957]0.976

0.9420 6 nO.97410.98710.998

0.979 0.971 0.952

0.987 0.972 0.942

61

Table 2.- SUPERSONIC ENGINE-FACE FRESSURE RFCOVERY DATA, pt2/p - Concluded

B l e e d exit setting B

tW

n TlTBE NO. I

.987 10.992

RAKE TUBE NO. NO.

- 1 2 1 3 - J - 4 -

1 '0 .963 0.98910.9981----

3 0 . 9 3 2 0 . 9 2 d 0 . 9 2 4 [ 0 . 9 0 6

5 -0 .926 0 .91610. .921]0.910 ~~

-5 ! 6. 0 . 9 8 6 ] 0 . 9 1 0

0 . 9 2 1 ] 0 . 9 0 4

0.91010.893 ~.

2 n0.968 l0 .962

4 10.90410.899

6 10.96110.959

TUBE NO.

5 0.950]0.970 10.955

0 . 9 0 9 [ 0 . 9 1 1 ] 0 . 9 0 8

0.96410.970 10.953

1 "6 -I

0.9ooJ

0 . 8 6 7 1

0 .8991

= 1.55 a = 5.0' mo/mw = "" mbp/m, = 0 .03

P ~ $ P ~ , =0 .931 q , l / m w = 0.113 apt2 = 0.121 P2/Pm = 3.8

RAKE TUBE NO. TUBE NO. RAKE NO. 2 1 NO. 1

w "" - .~ .~ .- 3-.! 21_5.1"6 -22" 1 0 . 9 5 1 , 0 . 9 6 1 0 . 9 9 8 - 2 0.999 ---- 0.990 0.976 0.960 : 3 0.886 ~~ [0 .899[0 .899 ~ 10.900-10.903 0.895 0.915 - - - - 4 0 . 9 1 1 " -

0.889 0.909 0 .912 ,0 .910 5 0.94510.946 10.953 b .95810.962 ~ ." 1 *0.954 0 . 9 1 1 , 0 . 9 1 4 6 0.896 0 .902 0 .897 10.895

62

"".._.. . .. . ,

Tab le 3.- TRANSONIC ENGINE-FACE PRESSURE RECOVERY DATA, ptdpt,

b = 0.6 a = 0.0' mi/m, = 0.329 CDa = 0.065

Pt2/Pt, I 0.952 ( x / R ) ~ ~ ~ = 4.180 Apt. = 0.075 ex i t s e t t i ng= Open Bleed -

TUBE NO. I TUBE NO.

1 1 2 4 3 1

0.973 0.967 0.951 -0.943 [ 3 0.970 0.965 0.956 .0 .947

I n 0 . 9 4 6 ~ 0 . 9 5 1 ~ 0 . 9 6 4 1 0 . 9 7 3

5 0.965

0.966

0.965

- 0.918

0.906 0.957 0.974 0.967 0.957 6 10.949 0.9141

0.926 0.968 0.974 0.963 0.951 ,0.944 4 0.912

0.888 0.959 0.977 0.94910.95410.965 2

TUBE NO. No- 1 1

2 n0.981

4 10.980

6 10.978

. -

TUBE NO.

2 I T 4 5 6 4

0.991 0.996 1.001 0.994 0.940

0 . 9 9 0 ] 0 . 9 9 ~ ~ ~ 0 0 1 0.998 0.982

0.985 10.992 70.996 0.992 0.966

3 1 4 1 5 6 0.976

0.967

0.976

0.99310.999 10.999

0.99110.996 10.994

0.993 I 0 997 11" 0 999

RAKE II TUBE NO. 1- II TUBE NO. I NO. j T

10.957 - ~~-

3 10.953 10.958

5 [0.951=

6 ' NO. 1 1 2 6 I 5 4 3

0.922 0.883 0.959 0.993 0.980 0.963 0.952 2

4 I 5 KiGkzii

3 0.972 ~____

0.973 0.912 4 10.948 10.958 10.973 0.907 0.979 0.992 0.966 0.983 .951

0.931 3.980 0.989

0.977

0.6 a = 0' mi/mw = 0.283 M, = = 0.106 Bleed

l m n TUBE NO. -

RAKE TUBE NO.

5 1 6 6 5 4 3 2 1

l.OOOl 0.9771 2 ~~

0.985 1.000 1.000 0.999 0.989 0.983 0.99810.9691 4 0.943- 0.994 1.000 1.000 0.995 .0 .986

1.00010.976] 6 0.965 0.996 1.000 1.001 0.992 0.985

I NO. ii 1 2 1 3 1 4

I 1 10.9831 0.9911 0.9991 1.000

1 3 10.9841 0.9911 0.99d 1.000

[ 5 10.9851 0.9901 0.9971 0.999

NO. ,I

1 2 1 3 1 4

10.9831 0.9911 0.9991 1.000

10.9841 0.9911 0.99d 1.000

10.9851 0.990] 0.9971 0.999

63

Table 3.- TRANSONIC ENGINE-FACE PRESSURE RECOVERY DATA, p t , /P t , -Con t inued

& = 0.6 a = 0.0' . mi/% = 0.339 cDa = 0.098

- pt2/pt, = 0.966 (x /R) l ip= 3.880 A p t 2 = 0.091 Bleed e x i t s e t t i n g = Open

TUBE NO. m TUBE NO. NO.

-1 0.956 0.975 0.99110.996 -2 1 -~ 3" 1 4

3 0.943 0.95810.p78]0.990

5 -0.957 0.966]0,98810.997

~. 4- 1 5 I. 6 0.99710.95810.885

3 0.995

0.988

0.988

=.I- 11 . 0.95810.982

0.95810.96i

0.94910.965

0.980

0.987 ~~.

0.99610.98310.929

0.994IO.98210.909 0.979 .

0.283 CD;, .= 0.143

~ . .

NO.

4 ! . . 5 1 6 1 I y I v 1.00110.99910.9781 2 . 10.98710.997

~~ 1.ooo~0.999~0.971[ 4 10.98510.991

__ 0.999[1.000~0.977~ 6 10.98810.993

TUBE NO. 1 -3 ! ! 5 . 1 . 6 - 1 1 . 0 0 0 ~ 1 . 0 0 0 ~ 0 . 9 9 3 ~ 0 . 9 4 4 ~

0.99911.000 [1.00010.989]

1 . 0 0 1 ] 1 . 0 0 1 [ 0 . 9 9 4 ] 0 . 9 6 6 ]

RAKE TUBE

&I= 0.6 a = 2.0' mi/m, = 0.329 CDa = "

Bleed e x i t s e t t i n g = ooen

". .. n TUBE N O .

0.6 a = 2.0' q / m , = 0.287 - "

Apt2 = 0.035 e x i t s e t t i n g s Open

M, = " 'Da - Bleed

= 0.990 4.180 (x/R), ip=

I II TUBE NO. I1 RAKE I I TUBE NO. ~- . "~

- 4 . I 5- " 1.00110.999

" 1.000~1.000 .".

-

-

0.9451

0.9841

0.9661

3 1.000

0.994

1.001

~~ -

__ "

-

64

T a b l e 3 . - TRANSONIC ENGINE-FACE PRESSURF: RECOVERY DATA, p t a p t , - Continued

RAKE NO.

1

3 5

"

RAKE NO.

1

3 5

~

NO.

4 1.000

1.000

1.000

E 11.ooo lo. 998

10.999

II u 1 no.987 10.982

10.989

TUBE NO.

! 2 ! 3 1 4 1 5 ! 6 q 0.995

0.988

0.996 ~ It" 1.000 11.001 10.999 0.946

0.998 1.001 1.000 0.984

1.001 1.001 0.995 0.968

I 3 ~0 .93710 .937~0 .93810 .94410 .93710 .9041 4 10.92910.928 0.893 0.958 Oi980 0.978 I 5 [ 0 . 9 3 9 [ 0 . 9 4 1 ~ 0 . 9 4 2 ~ 0 . 9 4 7 [ 0 . 9 4 ~ [ 0.9121 6 10.967 10.968

0.910 0.933 0.941 0.936

0.6 a = 5.0' mi/m, = 0.287 'Da - - M, = "

B l e e d

TUBE NO.

0.9691

0.9601

3.9690

RAKE

No. 2 10.987

4 10.966

6 10.988 __

2

0.998

0.967

d

TUBE NO. 1 3 1 4 1 I 5 1 1 6

65

II I

Table 3.- TRANSONIC ENGINE-FACE PRESSURE RECOVERY DATA, pt2/ptcx, - Continued

RAKE NO.

TUBE NO.

1 3 2 c 5 1 6 4

1

0.9171 6 0.948 0.952 0.946 0.947 0.955 5 0.910] 4 0.946 0.949 0.947 0.946 10.953 3 0.909 0.991 1.000 1.000 0.998 0.962

~~

I1 TUBE NO.

Mal= 0.6 a = 5.0' mi /m, = 0 .283 CDa = -- Bleed -

PtJPt, = 0.986 (X/R)l ip= 4 .030 Apt2 = 0 . 0 4 1 e x i t s e t t i n g = open

RAKEn TUBE NO. n TUBE NO. 1 4

.975

.922

.978

0.6 a = 8.0' CDa - -- - M, =

Bleed

..

1 ! 5. !.?A

97 10.985]----

6410.9631 0.9561

98]0.987]0.9521

66

I

Table 3 .- TRANSONIC ENGINE-FACE PRESSURE RECOVERY DATA, ptz/pt, - Continued

M, = 0.6 a = 8.0' mi/m, = 0.336 'Da - - -- Bleed - Pt2/Ptm= 0.949 (x/R)lip= 4.030 Apt2 = 0.123 exi t s e t t i n g = Open

I n n 1 pi? 1 3 I 5

TUBE NO. 1 2 1 3 1 4 1 5

I I I 1

TUBE NO. I 2 1 3 1 4 1 5 1 6

I E E I ! . TUBE NO. p - 1 1 I 2

I n0.981~0.973 0.972]0.971]0.967 0.971~0.97310.970

M , = 0.8 a = o.oo rni/m, = 0.332 CD,. = 0.074

&2/Pt, '0.912 (x/R)lip= 4.180 Apt2 = 0.122 ex i t s e t t i ng = Open

~~ ~~

Bleed

TUBE NO. 1

0.92610.92910.93810.903 0.92610.930]0.939[0.91?

6 0.846

0.837 0.842

L~

0.93310.935 10.939 10.904 I---- I 0.924 0.936 0.943 0.914 0.851

0.92710.93210.940 0.904 0.832

0.8 a = o.oo m i / % = 0.295 M, = CD,. = 0.098 ~.

Bleed

RAKE NO.

1

3 5

1.1 0.96710.979 0.97210.978 0.97110.985

TUBE NO. " -

. .. 3 I -4-T" 5 0.9931 0.99910.996 0.9861 0.990 [ 0-..984 0.9911 0.98510i983

6 0.957 0.944

0.946

II TUBE NO. I

67

Table 3 . - TRANSONIC ENGINE-FACE PRESSURE RECOVERY DATA, p t d p t , - Continued

TUBE NO. 6

0 .861

0.844

0.848

... . . . NO.

2

4 6

" " -

TUBE NO. 1

10.955 10.969 10.924 I---- I 10.950 10.966 10.949 10.865 1 10.96310.96810.93010.8361

TUBE NO.

3 4 5 0.999 1.000 0.999

2

0.992

0.986

0.989

TUBE NO. 1 ~~ 1

0.998 1.000 10.991]---- 1 3 . 1 5 . 1 6 - 1

~ 1 . 0 0 0 ~ 1 . 0 0 0 ~ 0 . 9 9 5 ~ 0 . 9 4 5 ] 10.999 1.000 ]0.998[0.958]

mh TUBE NO. 1 NO.

3 ! 4"! 5 ! 7 2 n0.91210.938[0.96810.988 10.925[---- ] 4 ~ 0 . 9 1 0 ] 0 . 9 2 9 ] 0 . 9 6 1 1 0 . 9 8 7 10.96510.8591

6 n0.91110.929 10.971 10.984 10.938 10.835 ]

0.8 ~"

a = o.oo mi/ '% = 0 . 2 9 1 M, = c~~ = 0.167

Bleed Ft2/pt,= 0.985 ( X / R ) l i p = 3.880 = 0 . 0 4 3 e x i t s e t t i n g = Open

RAKE II TUBE NO. 4

1.000

1.000

0.997

_ _ _

"

~

.

.1.1 .98010,993 [ O

.97610.987.]1

.981[O,988 [I

TUBE NO. ~~ "

3 1 4 1

68

Table 3 . - TRANSONIC ENGINE-FACE PRESSURE RECOVERY DATA, ptdpt, - Continued

% = 0.8 a = 2.0° mi/m, = 0.332 CDa - -- -

B l e e d - Pt2/Pt, =Os909 (x /R) l ip= 4.180 Apt* = 0 . 1 4 4 ex i t s e t t i ng= Open

RAKE NO.

1

3 5

II TUBE NO. RAKE II TUBE NO.

1 6 5 4 2 3 . N o d 1 2 I 3 I 4 I 5 I 6 ' 0 . 9 3 7 ~ 0 . 9 4 3 ~ 0 . 9 5 5 ~ 0 . 9 6 4 ~ 0 . 9 3 8 ~ 0 . ~ 0 . 9 3 6

0.847 0 .897 0.915 0 .909 0 .909 0 .909 4 0 . 8 4 3 - 0 . 8 9 3 ~ 0 . ~ ~ ~ 0 1 0 . 9 1 9 1 0 . 9 1 9

---- 0.909 0.949 0 .945

no.91310.91510.916To.920 0 . 8 3 3 ' 0 . 9 0 8 0.948 0 .939 0.932 0.929 6 0 .843 0 .902

I No- [-I] 2 I 3 I 1 0 . 9 6 9 1 0 . 9 8 4 ] 0 . 9 9 9

I ~ _ _ ~ o . 9 7 2 ~ o . 9 7 2 ~ o . 9 7 4

I 5 10.975 10.980 10.981

- I m j 1.000 0.987 ----

i 0.980 0 .977 0 .958

1.000 0.993 0 .940

1 r . L 1 * 1 3 1 4 I 5 6 TUBE NO.

1- 1 110.939 10.962 10.985 10.995

0 . 8 6 1 0 .922 0 .940 . 9 1 1 0 . 9 1 1 0 . 9 2 8

0 . 8 5 1 0 . 9 2 1 I 3 ~0.91310.91310.922]0.937

0 . 8 4 2 0 .960 ~ -~

RAKE II TUBE NO. I No+ I I 1 2 I 1 3 I I 4 1 1 5 I 1 6 I

2 b . 9 3 3 10.946 10.962 10.973 b.918

0 . 8 3 3 0 . 9 3 1 0.975 0.944 10.967 6 10.933

0 .866 0 .922 0 .933 0 .91210.926 4 10.907

----

0.8 a = 2.0° m i / % = 0.292 cDa - M, = - --

B l e e d

TUBE NO. "I W E NO.

4 ! 5 1 ~ - 6 . No* 11 1 I 2 3 4 5 6

1 . 0 0 0 ~ 0 . 9 9 8 ~ 0 . 9 5 7 ~ \0.98110.995 ~~ o.999 l.ooo 0.996 ---- 1 . 0 0 0 1 0 . 9 9 2 [ 0 . 9 4 5 ~ 4 10.966 0 .976 0 .993 1.000 0 .999 0 .970

0 . 9 9 8 1 0 . 9 9 3 ] 0 . 9 5 3 0 6 10.980 0 .993 1.000 1.000 0.995 0.945

69

Table 3 . - TRANSONIC ENGINE-FACE PRESSURE RECOVERY DATA, p t d p t , - Continued

n m n TUBE NO.

I - I: .

U T 10.948

10.885

10.944

0.956[0.8991----

0.888 - IO. 871 10.836

0.95210.90010.820 ". -

RAKE TUBE NO. NO. 1 2 3 1 4

~~

1 ~0.975 0 .983 0 .995[1 .000

3 0.973 0.964 0.96370.963

1 5 1 6 lo. 998 l o . 950

10.95310.937

~ 10.95310.938

: %qm:""I'"* ! 5 . I .5 -1 1 i /0.g82 ~ 0 . g g 7 ~ 0 . g g 7 10.983 I---- I

4 10.951 10.952 10.956 10.959 10.950 10.9421

6 10.984 10.997 10.999 lO.999 10.986 10.932 1

IRAKE " "

I . 5 1 9, i0.96g10.84? ]0.891[0.851

10.89810.855 ~ .. .

n

10.967 IO. 968

TUBE NO. I

0.8 M, = a = - 5.0' mi/m, = 0.292 cDa -

- "

Bleed

70

TUBE NO. I

I

Table 3.- TRANSONIC ENGINE-FACE PRFSSURE RECOVERY DATA, pt-pt, - Continued

% = 0.8 a = 8. 0' mi/m, = 0.332 cDa - - "

- Bleed pt2/pt, =0.897 (X/R)lip= 4.180 Apt2 = 0.205 exit setting= Open

RAKE NO.

1

3 5

- "

I TUBE NO. TUBE NO. W I 1 6 5 4 3 2 1 No. 6 5 4 3 2

-0.983 ---- 0.882 0.933 0.928 0.940 -0.964 2 0.825, 0.957 0.999 1.000 0.998

~0.889]0.88710.88210.882]0.863~0.825~ 4 [0.88010.872 0.816 0.887 0.930 j0.89310.88410.88210.87610.861~827~ 6 10.96710.970.933

0.819 0.848 0.866 0.876

TUBE NO.

3 1 4 i

.000~1.000

.94010.938

.942 I O . 936

TUBE NO. I

%a= 0.8 a = 8.0' w/m, = 0.338 = "

Bleed Ft2/pt, = 0.909 (x/R)lip= 4.030 Apt* = 0.198 exit setting= Open

RAKE NO.

1

3 5

II TUBE NO. TUBE NO. I II TUBE NO. 1 1 2 3 4 5 6

7

0.977 0.978 0.975 0.971 0.902 ---- 0.878 0.869 0.872 0.872 0.859 0.836-

10.91310.89510.884 0.875 0.859 0.8311 6 -'0.981 0.978 0.976 0.976 0.916 0.820

0.999~1.000~1.000 0.966 0.833 2 0.977 0.978 0.975 10.971 0.902 ----

0.90010.887 0.884 0.863 0.830 4 0.878 0.869 0.872 0.872 0.859 0.836-

0.89510.884 0.875 0.859 0.8311 6 "0.981 0.978 0.976 0.976 0.916 0.820

0.8 a = 8.0' mi/m, = 0.292 M, = ~

- 'Da - "

Bleed

1-n TUBE NO. I NO. 1 1 2 ~_______

1 3 1 4 1 I ~ . ; ~ J 0 . 9 9 6 ~ ~ ~ ~ 0 0 0 ~ l.ooo [ 3 00.9701 0.9551 0.9491 0.949

0.970 0.945 0.949 0.953

m I TUBE NO.

5 I 6 6 5 4 3 2 1 No*

0.998]0.940-

0.929 0.987 0.9471 0.9341 6 ~ 0 . 9 9 2 ~ 0 . 9 9 8 ~ 1 . 0 0 0 ~ 1 . 0 0 0

0.935> 0.939 0.942 0.941 0.94210.9310 4 10.94010.941

---- 0.982 1.000 0.998 1.000 >0.991 2

71

I I I m.., . ,.,.,,. .-.,. , , . . ... .. .. ... . -. ~

Table 3. - !TRANSONIC ENGINE-FACE PRESSURE RECOVERY DATA, p t a p t , - Continued

Bleed e x i t s e t t i n g = Open

kF ~

l o . 892 10.904

10.88710.903

n0.890 10.889

TUBE NO. TUBE NO. -

3. -1 !L: 0.9J210.914

1 "

"

5

0.888

0.881

~. " -

0,884

3'

~

0.919

0.905

0.927

4 I 5 l"6 I 1

0.932 10.880 I----

0.93410.885]0.798

0.91410.892 10.810 1 10.89210.897

0.889 0.899

0.90810.917

0.~9-16 10.92J

a = o M, = 0.9 -

' t z - - e x i t s e t t i n g = open __

TUBE NO. 1 7 ~ 3 ! 1 5 . 1 6 - 98110.985 10.990 10.982 I---- 9 8 3 ~ 0 . 9 9 8 ~ ~ . 0 0 0 ~ 0 . 9 9 ~ ~ o . ~ 4 6

97710.987 10.994 10.990 10.937

0.067

1 6 1

,935 I .955

.941]

]0 .971]0. I-" 10.965 I O .

10.969 10.

RAKE TUBE NO. NO. 1 2

~~

3 1~~ 4 1 0.964 0.974 0.99470.999

3 0 . 9 7 0 0 . 9 7 8 0.986 0.991 5 ,0.968 0.978 0.989,0.994

RAKE NO.

2

4 6

TUBE NO.

0.897 0.902

3 0.935

0.928

0.936

___ . ".

. ~ _ _ .

10.929 10.947

10.937 . . 10.946

0.9 a = 0.0' mi/m, = 0.298 C D ~ = 0.144 M, = Bleed

W t m = 0.982 (x/R)lip= 4.030 Apt* = 0.069 e x i t s e t t i n g = Open

RAKE NO. - 1

TUBE NO. I1 I1

-. 0 999 1 1.000 , 1 0.999 1 0.953 I 1 ;0.977;0.991

0 9 9 8 1 0 4 0 . 9 9 8 0.940 0.977 0.988

0.9991 1.000 0.996 0.932 4 j0 .97310.984 3 5

72

T a b l e 3 . - TRANSONIC ENGINE-FACE PRESSURE RECOVERY DATA, pt-pt, - C o n t i n u e d

M, = 0.9 a = o.oo mi/m, = 0.336 cDa

= 0.152

TUBE NO. RAKE TUBE NO. I I NO. I 1

I 3 ~ 0 . 8 8 0 ~ 0 . 9 0 5 ~ 0 . 9 3 8 ~ 0 . 9 6 4 ~ 0 . 9 1 8 ~ 0 . 8 1 2 ~ 4 10.884 0.903 0.943 0.967 0.945 0.831

0.822 0.910 0.947 0.970 0.915 ---- 0.946 0.905 0.939 0.971 -0.879

I 5 10.8831 0 . 9 1 1 ~ 0 . 9 4 3 1 0 . 9 6 8 ~ 0 . 9 2 6 1 0 . 8 0 6 ~ 6 10.88910.91710.956 0.970 0.911 0.804

1 1 2 1 3 1 4 5 - 6 [ 1 2 3 4 5 6 '

NO. 1

RAKE NO.

1

5- 5

n TUBE u ." 1 1 m gO.97410.98710.9991

10.97310.98610.9991

nO.97810.98310.9961

TUBE NO. 1

4 ~ 0 . 9 7 4 ~ 0 . 9 8 6 ~ 1 . 0 0 0 ~ 1 . 0 0 0 10.997 0.948

6 ~ 0 . 9 7 8 ~ 0 . 9 8 7 ~ 1 . 0 0 0 ~ 1 . 0 0 0 10.996 0.942

TUBE NO. TUBE NO. I N o * ~ 1 1 2 1 3 1 4 1 5 1 1 2 1 3 I 4 I 5 1 6 1 ]0.816[0.83510.8621

3 ~0 .81510.827[0 .850]

5 ]0.81910.840]0.8661

0.88510.86310.7711 2 10.828 10.852 10.877 10.895 10.852 0.768 0,87110.84810.7591 4 10.818 10.83310.861 (0.884 (0.862

----

0.883 0.857 0.760 6 0.824 0.844 0.871 0.894

1.0 a = o.oo mi/m, = 0.299 M, = 'Da = 0.145

B l e e d

P t 2 / P t r n = 0.960 ( x / R ) l i p = 4.180 A P t 2 = 0.069 e x i t s e t t i n g = O D e n

IRAKEI TUBE NO.

I 1 10.9561 0.9641 0.9791 0.98510.9841 0.940 [ 3 10.9551 0.9601 0.9691 0.97010.9631 0.920

I " 5 n0.95710.961] 0.9651 0.9661 0.9711 0.926

RAKE NO.

2

4 6

_____

____

TUBE NO. I

7 3

T a b l e 3 . - TRANSONIC ENGINE-FACE PRESSURE RECOVERY DATA, ptJpt, - C o n t i n u e d

r n

RAKE NO.

2

4 6

." - -

TUBE NO. RAKE TUBE NO. NO. "3" 1 2 4 ~~ " 1 5 ! 6 . .

0.888 0.897 0.916 0.91910.882]0.794

3 0.889 0.898 0.913]0.915[0.864]0.775

5 0.894 0.902 0.911]0.91210.86810.784

"

~

2

901

890

902

~~ 4 1 5

0.924l0.865

0.90710.867

0.92110.863

3 0.913

0.900

0.917

_ _ _

% = 1.0 a = o.oo mi/m, = 0.303

l m I No* i 2

4

T 0.999]1.000

1.ooo~1.ooO

TUBE NO.

0.996 11.000

1 TUBE NO. I r k). lo. lo

7161 990 I"" '1

NO. 1 2 3 4 1 5 1

1 .0.971 0.983 0.998 0.999)0.9981

6 0.945

0.927

0.934

~~

10.975 10.990

10.970 10.983 1

3 0.972 0.983 0.999 0.999 0.993

5 0.975 0.984 0.997 0.994 0.997 -~~

6

RAKE NO. I 5 1 6

10.889 I---- 10.914 10.800

10.885 I O . 774

. . . I

- " 1

TLE3E NO.

-3 I TqiTg 0.91110.931

0.91110.934

-

1 5 1 6 F 9 1 5 ] 0 . 8 0 0

10.89110.783

10.89410.787

F io. 882

1 3 10.914

10.918

10.930

1 4 - io. 942

Io. 945

10.948

10.859

00.858

no. 862 ~...

2 4 6

0.861 0.876

1.0 a = 0.0' m i b , = 0.302 M, = = 0.240

B l e e d

5 t 2 / P t , = 0.980 (x /RJ lLp= 3.880 A = 0.074 e x i t s e t t i n g = Open

P t 2 ~-

1 .~

1 5. 1.00010.987

1.000]0.996

1,00110.996

6 . - ""

0.943

0.932

74

I

Table 3 . - TRANSONIC ENGINE-FACE PRESSURE RECOVERY DATA, p t2 /p t , - Cont inued

1

0.839

0.810

0.812

2

0.862

0.825

0.825

2

0.970

0.962

0.959

T U B 1

3 0.979

0.971

0.958

NO.

0.975 10.95810.914

0.95810.95110.916

RAKE NO.

2

4 6

II TUBE NO. I

% = 1.0 a = ~~ 2.0° mi/%, = 0.333 cDa = "

Bleed FtJPt , =0 .872 (x /R) l ip= 4 .030 Ap t2 = 0.214 e x i t s e t t i n g = Open

TUBE NO. NO.

1

3 5

10.877 10.875

10.883 10.887

3 0.947

,0.887

10.895

3 0.947

0.887

10.895 0.897 0.858 0.778

RAKE! TUBE NO. NO. 1 1 1 ~ 2 1 3 6 4 5 - ~~

2 10.902 10.9~10

0.769 0.932 .874 0.924 0.899 0.907

0.788 0.868 b.848 0.864 4 10.85610.858

---- 1 0.925 b .861 0.917

M, = 1.0 a = 2.0° CDa - -- -

Bleed

Pt JPt, = 0.980 (x/R)l ip= 4 .030 A = 0.079 e x i t s e t t i n g = Open P t 8

RAKE NO.

2

4 6

10.978l0.993

n0.96210.974

TUBE NO. I 3

0.997

0.994

1.001

1.000 0.995 ---- 1.000 0.996 0.947

1.000 0.996 0.932

75

Table 3. - TRANSONIC ENGINE-FACE PRESSURE RECOVERY DATA, p t , / p t , - Continued

TUBE NO. 1 TUBE NO. . ..

4 I 1 - 6 i 0.875 10.821 I---- I 0 .80610 .784[0 .745]

0.88610.828 10.738 1

T -2 lo . 870

lo. 810

10,883

6

746

7 4 4

749

-. .

Bleed - P t 2/Ptm = 0.949 (x/R)lip= 4.180 Apt, = 0-087 e x i t s e t t i n g = Open

TUBE NO. 1 M E TUBE NO.

1 3 10.978

Io. 984

10.933

I

1 5 . ! .6 -I 3.986 10.969 I---- 1 3.934 10.929 10.915 ] 3.988 p .975 10.910

- 1 4 I - 5 10.985 10.985

10.933 10.924

10.93110.930

% = 1.0 a = 5.0' mi/m, = 0.333 % = " Bleed

0.277 e x i t s e t t i n g = Open ~

.~ "~ n W E NO. 1 RAKE II TUBE NO. FAKE NO.

2

4 6

~ .

1.0 a = 5.0' mi/mm = 0.303 CDa - -- - M, =

Bleed

76

Table 3. - TRANSONIC ENGINE-FACE PRESSURE RECOVERY DATA, ptJptw - Continued

TUBE NO.

1 1 2 1 3 1 4 1 5

I 1 _0.93910.96810.98910.98510.879

I 3 10.79610.78910.78710.78610.773

I 5 nO.79910.79710.79910.79910.783

TUBE NO. I

0.7191 2 10-.892 10.854 10.846 10.852 10.799 ---- 0..7291 4 10.768 10.76910.774 10.775 0.760 0.730

0.7380 6 10.890 10.848 10.848 10.851 0.811 0.719

RAKE NO.

1

3 5 .. . .

n TITBE NO. RAKE II TUBE NO. I 3

-980

.9 13

.983

% = 1 . 0 a = 8.0' 0.333 -

cDa - "

Bleed PtJPt, 0.850 (x/R)lip= 4.030 npt2 = 0 . 3 0 4 e x i t s e t t i n g = Open

W E NO. TUBE NO. I

1.0 a = 8,O' m i / % = 0.303 CD, - " M, = -

Bleed = 0 .955 (x /R) l ip= 4 .030 Ap t2 = 0 . 0 9 4 e x i t s e t t i n g = Open

.. I RAKE II TUBE 1\10. II II W E NO. I

3 1 4 1 5 1 6 1 I I

77

Table 3 - - TRANSONIC ENGINE-FACE PRESSURF: RECOVERY DATA, pt;/pt ,- Continued

M, = 1.1 a = o.oo m i b , = 0 . 3 2 1 -

P t JPt, = Apt2 -

'Da - 0 . 1 7 8

0 .812 ( x / ~ ) ~ ~ ~ = 4.180 Bleed - - 0.187 e x i t s e t t i n g = Ope*

TUBE NO. ~ .~

I

1 ---- 0.829 0 .881 0 .861 0 .828 0 .799 2 0.753 0 .835 0 .858 0 .836 0 .800 0.785

3 0.727 0 .836 0.869 0 .845 0 .808 0 .790 6 0.7361 0 . 8 3 1 0 .858 0 .845 0 .804 0 .771 5 0.755 0.845 0.858 0 .833 0.807 0 .781 4 0.733 0.832 0 .870 0 .838 0 .802 0 .786

M, = 1.1 a = o.oo m i / m , = 0.333 CD,. = 0 . 1 8 6

Bleed &$Pt, = 0.853 (X/R)lip= 4.030 Apt2 = 0 . 1 7 8 e x i t s e t t i n g = Open

1.1 a = o.oo m i / m , = 0 . 3 1 1 M, = %. = 0.220

Bleed e x i t s e t t i n g =A

78

LL! 0.8401

0.8361

0.8361

1 TUBE NO. RAKE

~

2 1 3 " l 4 I 5 I 6 N o . . 1 1 2

0 . 8 5 9 ~ 0 . 8 9 7 [ 0 . 9 3 2 [ 0 . 9 0 8 ~ 0 . 7 7 8 ~ 2 ] 0 . 8 3 8 ] 0 . 8 6 0 1 0.85210.89710.92510.88310.768l 4 0.839 0 .860

0 .85410.89210.91910.88510.764~ 6 ,0 .841 0 .878

TUBE NO.

3 4 5 6 0.890 0 .928 0 .871 ---- 0.908 0 .931 0 .898 0 .786

0.920 0 .937 0 .879 0 .764

Cn = 0 . 2 8 6

TUBE NO.

-fl .804 lo .797 l o . 8 0 2 [ 0

6

.718

.702

.703

1 2

io. 779

lo. 778

lo. 780

1 .2 a = o.oo mi/m, = 0.307 M, = cDa -~ = 0 . 1 6 1

B l e e d

RAXE NO.

1

3 5

n TUBE NO.

ii 1 . I =! I. . . 3 I - 4 - r v ~ 0 . 8 8 8 ~ 0 . 8 9 5 ~ 0 . 9 0 9 ~ 0 . 9 2 0 ~ 0 . 9 1 7 ~ 0 . 8 8 7

~ 0 . 8 9 2 ] 0 . 8 9 7 ~ 0 . 9 1 4 ~ 0 . 9 2 2 ~ 0 . 9 0 8 ~ 0 . 8 7 4

~ 0 . 8 9 7 ~ 0 . 9 0 3 ~ 0 . 9 1 5 ~ 0 . 9 1 7 ~ 0 . 9 1 6 ~ 0 . 8 7 5

RAKE NO.

2 4 6

.. 1 TUBE NO.

79

Table 3 . - TRANSONIC ENGINE-FACE PRESSURE RECOVERY DATA, p t a p t , - Continued

TUBE NO. . . " - . _ . ..

TUBE NO. 2

. .. ".

.9 66

.966

.964

- ._ .

1 .I 5".J.. . 6 . . I

10.964[----

10.96110.947l

10.968]0.940] _" .

I n I RAKE 11 TUBE NO.

"

I 5. 10.857

10 848

10.848

k = 1.2 a = o.oo ~.

E Io. 822

lo . 809

10.833 . "

0.317 -

0.085

II -

[O. 946 10.963

941 LO. 954

1 0 . 9 i d 0 . 9 5 9 " ..

.. .

.. . .. ,

TUBE NO. ~

3 ! 4.. 0.852 10.880

0.843 I O . 868

0.87210.894

Da Bleed

e x i t s e t t i n g

= 0.176

I z.. I . I

'..974 I"" J

.985 1 0 . 9 1 q . . -

.985 IO. 926 1

80

Table 3 . - TRANSONIC ENGINE-FACE PRESSURE RECOVERY DATA, p t a p t , - Continued

% = 1.3 a = o.oo mi/m, = 0.324 - cDa = 0.137

- pt2/pt, = 0.745 (x/R)xip= 4*180 A p t 2 = 0 . 1 9 9 e x i t - s e t t i n g = Open Bleed

M, = 1.3 a = o.oo rni/rn, = 0.313

P t ./Pt, = 0.858 ( x / R ) l i p = 4 . 1 8 0 n p t 2 = 0.055

cDa = 0.141 Bleed - e x i t s e t t i n g = Open

I TUBE NO. 1 2 3 . - . 4 -- l . 5 6 /O:h,;z 0.873 0.87110.856 0.831

u q . 8 5 8 ~ 0 . 8 5 7 ~ 0 . 8 7 0 ~ 0 . 8 7 0 ~ 0 . 8 4 9 ~ 0 . 8 3 4

~ 0 . 8 6 7 ~ 0 . 8 7 0 ~ 0 . 8 7 5 ~ 0 . 8 7 2 ~ 0 . 8 6 1 ~ 0 . 8 3 2

TUBE NO. 1

.. !.- .899 10

. g o 8 10

. ..

.891[0

. . .

TUBE NO. . I "

I . 3 . .!. ! .909]0.910 10.

.91210.907). ~~

p~

.903[0.905 10.

516 901 i o . 882

896 10.878 901 I O . 875

1.3 a = o.oo q / m , = 0.341

= 0.802 (x/R)ll,p= 4.030 Pt2 = . .0 .203 ex i t s e t t i n g s Open

% = 0.146 M, =

Bleed

k 2 / P t m

I R A K E I I TUBE NO.

. - 6 1 5 1 6 ' . .

.872 IO. 8431 0.728

.858]0.82610.712

.85510.814] 0.716

2 1 0 . 8 0 7

0.8037

0.8111

W E NO. I 3

0.832 -

0.836

0.851

Table 3 .r TRANSONIC ENGINE-FACE PRESSURE RECOVERY DATA, p t 2 / P t m - Continued

% = 1 . 3 a = 0.0' - m&, = 0.328 CDa = 0.156

Bleed - pt2/pt,= 0.945 ( x / R ) l i p = 4.030 Apt2 = 0.086 exit s e t t i n g = Open

I 5 . 10.961

10.975

10.974

I I -6 1 I O . 859 I 10.9141

10.8991

RAKE TUBE NO. 1

"_ ~~ ~-3- 1

1 ,0.956 0.962 0.9771

3 '0 .959 0 .960 0 .9751

5 -0 .963 0 .966-0 .9741

NO.

4 1- 0.98110

0 . 9 7 9 1 o 0.97310

I -

TUBE NO. 1 2 3 " I 4 1 5 I. .__4 1 I

0.96910.97110.98110.9771---- 1 0.9621 0 .97510 .981 [0 .980 [0 .943 ]

0 .965 ]0 .97610 .982 ]0 .980 [0 .932 ]

Mm= 0.6 a = o.oo = 0 . 3 2 9 ' D a = 0 . 0 6 5

Ft$pt, = 0.956 ( x / R ) ~ ~ ~ = 4.180 Bleed

A p t 2 = 0.094 e x i t s e t t i n g = Closed

82

Table 3.- TRANSONIC ENGINE-FACE PRESSURE RECOVERY DATA, ptz/Pt, - Continued

RAKE NO.

1 3 5

TUBE NO.

. 0.96810.97410.973 ..

0.971[0-. 9 7 4 G 3

- 0.968b.97310.975 -

RAKE II TUBE NO. I NO. ~. "

2

4 6

"

TUBE NO. I

83

Table 3 .- TRANSONIC ENGINE-FACE PRESSURE RECOVERY DATA, p t d p t , - Continued

M, = 0.6 Q: = 2.0° mi/m, = 0.329 - cDa -

ex i t s e t t i ng = Closed

"

Bleed - pt2/pt,= 0.955 ( x / R ) ~ ~ ~ = 4.180 = 0.098

I n

I RAKE II TUBE NO. 5 .977

L.L -

.965

.965 __

M, = .0.6 a = 2.0°

. .

TUBE NO. I m 0.98410.992

0.95410.967

0~. 981 [ 0.991,

l0.95510.879 I 10.969 10.9241

]0.969]0.9001

CDa = "

Bleed

M, = 0.6 a = 2.0° d m , =

. __ .~ . - --

TUBE NO.

0.966 0.981 0.99110.999 0.981 0.9101 2

0.952 0.958[0.97510.98710.966[0.9041 4 0.951 0.959~0.974~o.985~0.969~0.907~ 6

0.336 . ~ CDa = "

Bleed 0.099 exi t se t t ing = Closed

-h TUBE NO. I 3 ! .?j 5 1 . 6 1

i0.967 ]O.981]0.990 10.993 10.958 ]0.881]

~0.941~0.946~0.962~0.980 10.97410.925 I 10.963 " 10.97510.989 10.995 .. - 10.969 10.9041

.. " - ." . . "" --

RAKE II TUBE NO. "

1

.~ .968[0

.9701 o

6 .

.960 ~- -

,953 ~~

.960 I .-

m NO.

2

4 6

. ~

["& 1[O.966LO2967

10.95510.957

[0296310.968

.. .

TUBE NO. 1 ~~ . .

-3 i---i 0.97110.977

0.97310.977

" ~

0.96310.966

5. I. '. 972 10 .971[0

.965 10

84

I

T a b l e 3 . - TRANSONIC ENGINE-FACE PRESSURE RECOVERY DATA, p t2 /p tm - Continued

I NO.

1 3 1 5

M, = 0.6 a = 5.0' q / m , = 0.329 " 'Da =

- B l e e d pt2/ptm =0.951 ( x / R ) l i P = 4.180 A p t 2 = 0.114 exi t s e t t i ng= Closed

TUBE NO. RAKE TUBE NO.

3 6 5 4 0.981

0.989 10.980

0.920 0.942 0.941 0.939 0.875 0.958 0.989

0.892 0.964 0 . 9 4 1 ~ 0 . 9 4 2 ~ o . 9 5 0 j 0 . 9 5 8 j 0 . 9 5 4 j 0 . 9 0 7 f ~____ 6 10.963i0.970

RAKE NO.

1

3 -. "

I u "" 1

10. 955 IO. 928 IO. 930

TUBE NO. I W 0 n 1

no. 943

~

10.942 10.923

10.940

TUBE NO. I 3

0.942

0.926 0.942

~____

RAKE NO.

1 3 5

_ .

TUBE NO. RAKE TUBE NO. 1 2 2 3 4 5 6

n0.95210.95110.961[0.968 10.959 10.906 1 4 10.930 10.936 10.945 70.956 0.956 0.930

~0.95110.95210.96310.973 ]0.962[0.909 I 6 00.983 10.984]6994 10.997 3.961 0.893-

0.6 a = 5.0' mi/m, = 0.283 M, = 'Da - - "

B l e e d

P t J P t m = 0.960 ( x / R ) l i p = 4.030 = 0.033 ex i t s e t t i ng= Closed

RAKE NO. 1

0.964

0.953 0.953

NO. 1-n

0.97010.96510.953 2 ;0.966 0.971 0.96210.96110.947 4 .0.949 0.950

"_ 0.960~0.960~0.951 6 0.96810.968 -

85

Table 3 .- TRANSONIC ENGINE-FACE PRFSSURF: RECOVERY DATA, ptapt, - Continued

% = 0 . 8 a = o.oo mib, = 0.338 CD,. = 0.092

Bleed iSt$Pt, = 0.934 (x/R)lip= 4.030 A p t 2 = 0.169 exit setting = Closed

0.8 a = 0.0' mi/% = 0.292 M, = C D ~ = . 0.124

Bleed fj /p = 0.945 (X/R>li = 4.030 0.038

= exit setting = C l o s e d

86

"

% = 0 .8 a = o.oo mi/m, = 0.339 C D ~ = 0.123

P t JPt, Bleed - = 0.944 3.880 apt2 = 0.167 ex i t s e t t i ng= Closed

% = 0.8 a = o.oo mi/rn, = 0.291 C D ~ = 0.167 Bleed -

pt2/pt, = 0.969 ( X / R ) l i p = 3.880 npt2 = 0.066 ex i t s e t t i n g = Closed

NO. 1 2

1 n0.94710.962

3 n0.94610.961

5 10.94510.962

TUBE NO.

3 1 4 1 5 0.9811 0.99610.988

0.98310.996 10.983

0.9531 2

0.9401 4 0.9491 6

1

0.949

0.947

0.947

TUBE NO.

%o= 0.8 a = 2.0° d m , = 0.332 CDa = "

Bleed &2/Ptm = 0.918 ( x/RIlip= 4 .180 ap tT = 0.180 e x i t s e t t i ng = Closed

- N O - 1 - 1 1 2 I 3 ~ 0 . 9 0 2 [ 0 . 9 1 2 [

1 !0.925[0.948[

5 10.8'3810.9051

0.838 1 0.829 I 0.842 I

TUBE NO. I 2 1 3 1 4 1 5 1 6

I I I 1 I

! N o - ii I . ! 2 ! .3 I . .4 I 5 1 6 U N O .

I n o . 9 0 9 ~ 0.g141 0 .924 0 .g28~0 .g21~ 0.8g31 =!

I 3 10.8931 0.8981 0 .904 0.9101 0.9031 0.8851 4 I 5 10.8951 0.8991 0.905/ 0.9O710.9041 0.8881 6

0.295 cDa - - "

Bleed 0.050 e x i t s e t t i ng = Closed

I TLIBE NO. 1 II 1

- 1 2 1 3 4 5 6

10.9041 0.911!915.0.920 *O. 909 ,--

~~0.88910.89410.898LO.904 10.898 0.883

0.901 0.908 0.916 0.921 0.908 0.884

87

, . . . . ...

% = 0.8 a = 2.0' mi/m, = 0.338 CDa - "

- pt2/pt, = 0.932 ( X / R ) - J ~ ~ = 4.030 Apta = 0.180 . exi t s e t t i n g = Closed

-

Bleed

RAKE TUBE NO. . ~.

4. :.

.p96 10.960

~. .972]0.936 ~ ..

. 9 7 1 ] 0 . 9 2 4

... . .

TUBE NO. " .

I T~ ..." - .. . . "" I 4 - 1 . , 5 I 6 J 0.94610. -95910.978]0 .983]0 .917] - - - - I 0.905 b-:916 10.940 10.958 10.951 10.872 1 o . 9 4 3 ~ 0 . 9 5 5 ~ 0 . 9 8 0 ~ o ~ . 9 8 4 ~ 0 . 9 2 3 ~ o . 8 2 8 ~ .~

RAKE TUBE NO. NO. 1 2 3 14 I " 5

0.945 0 .951 0 .964 0 .97210.966

3 0.932 0.940 0.944 0.95570.952

5 0 .933 0 .936 0 .944 0 .951 0 .954 ~~

TUBE NO.

% = 0.8 . .. a = 5.0' - mJm, = 0.332 - 'Da - -- Bleed

Ft$Pt, = 0.912 (x /R) l ip= 4.180 Apt2 = 0.202 ex i t s e t t i ng = Closed

88

M, = 0 . 8 a = 5.0' mi/m, = 0.338 - -- cDa - Bleed - pt2/ptm = 0.929 ( x/R)lip= 4.030 Apt2 = 0.194 ex i t s e t t i n g = Closed

n TUBE NO. II RAKE II TUBE NO. I 2 1 3 1 4 1 5 1 6

I I I I 3 1 4 0.99911.000

0.93010.947

0.9341 0.948

1

0.965

0.920

0.917

2

0 .996

0.917

0.917

1

0.970

0 .880

0 .968

0 .976

0.917

0.927

0.8451 2

0.8360 4 0.8460 6

TUBE NO. 3 I - 4 5 6 -

.962 0.970 0.950 ----

.928 IO. 933 0 .930 0 .915

.95710.970 70.952 0 .911 ~~

W E NO. W E NO. I 1

0.883

0.882

0.883

2

0 . 8 9 1

0 .895

0.894

0 .90010.91110.864

0 . 9 0 7 ] 0 . 9 0 8 [ 0 . 8 5 8

-_ 0.911]0.91310.85210.7671 6 ~~0.87910.89110.912 10.909 10.849 10.761 I 1.0 a = o.oo mi/m, = 0.299 M, = 'Da

= 0.145 Bleed

Pt2/Pt, = 0 . 8 8 4 ( x / R ) ~ ~ ~ = 4.180 Apt2 = 0 . 0 5 1 e x i t s e t t i n g = Closed

TUBE NO. I TUBE NO. RAKE NO.

1

3 5

3 1 4 I 5 I f . ! 2 I.. ..3 ! 4 1 5 0.87910.8811 0.8931 O.gO3[0.896

2

0 .884

0 .879

0.886

1

0.879

0.876

0.880

0 .8621 2 -

0.8671

o . 8 6 0 n 4 0.8771 0.8821 0.8921 0.89410.8g2

0.876 0.897 0.900 0.897 0 .885

89

Table 3 . - TRANSONIC ENGINE-FACE PRESSURE RECOVERY DATA, ptz/pt, - Continued

M, = 1.0 a = o.oo q / m , = 0.333 C D ~ = 0.138

Bleed - P t JPt, = 0.908 ( x / R ) ~ ~ ~ = 4.030 apt2 = 0.199 e x i t s e t t i n g = Closed

k = 1.0 a = 0.0' mi/m, = 0.303 = 0.178 Bleed -

PtJPt, = 0 . 9 3 4 (X/R)lip' 4.030 Apt2 = 0.056 e x i t s e t t i n g = Closed - - -. - - " - . .

n i B E NO. TUBE NO. I NO. 1 2

, I I

1 ---- 0.894 0 .906 0 .897 0 .886 0 .883 2 0 .874 0 .911 0 .919 0.909 0 .896 0 .891

3 0.857 0.879 0 .882 0 .880 0 .865 0 .867 4 0.857 0 .883 0.892 0 .885 0 .869 0 .871

5 0.857 3.896 0 0 . 9 0 5 0 .889 Lo.884 6 0.859 0 .888 0 .887 0 .888 0 .871 0 .878 __________

1 . 0 a = 2.0° d m , = 0 . 3 0 3 'D, - " M, = -

Bleed

@tm = 0 . 9 3 1 (x/R)lip= 4.030 Apt* = 0.071 e x i t s e t t i n g s Closed

90

Table 3.- TRANSONIC ENGINE-FACE PRESSURE RECOVERY DAW, p t z / p t * - Continued

0.299 cDa - - -- Bleed

0.114 e x i t s e t t i n g = C l o s e d

TUBE NO. NO. 1 ! ! 3.. ! 4 T... 5 1 ~ 0 ~ ~ 0 6 ~ 0 . 9 1 5 ~ 0 . 9 3 0 ~ 0.93810.92110.871n 2 3 ~ 0 . 8 5 6 1 0 . 8 5 8 1 0 . 8 6 3 1 0 . 8 6 9 ] 0 . 8 6 2 ~ 0 . 8 4 3 ~ 4 5 u0.85610.863]0 .86810.865]0 .86910.8491 6

TUBE NO.

1 2 1 3 1 4 1 5 1 6 .88410.880

.85110.855

.887 10.885

0.886

0.862

0.892

0.897I0.8811---- I ! I 1

0.86510.85610.8381 I I I

0.893 0 .887 0 .851

N o - U ~ 1 2 1 3 ~ 4 1 5 1 2

I 1 [ 0 . 9 1 3 ~ 0 . 9 0 9 ~ 0 . 9 1 3 [ 0 . 9 1 2 ~ 0 . 9 0 5 ~ 0 . 8 9 9 ~ -2 )0 .93510.948

[ 3 ~ 0 . 9 0 6 ~ 0 . 9 0 6 ] 0 . 9 1 4 ~ 0 . 9 2 2 ~ 0 . 9 1 7 [ 0 . 8 9 0 ~ 4 10.898101902

[ 5 ~ 0 . 9 0 9 ~ 0 . 9 0 8 ~ 0 . 9 1 6 ~ 0 . 9 2 1 ~ 0 . 9 2 4 ~ 0 . 8 9 3 ~ 6 00.93710.948

TUBE NO. ~ _. .

- N o . u l 1 1 3 J I 5 1, ~ 0 . 7 6 7 ~ 0 . 8 0 1 [ 0 . 8 4 2 ~ 0 . 8 6 6 ] 0 . 8 3 1 [ 0 . 7 2 7 ~

3 [ 0 . 7 6 9 [ 0 . 7 8 8 [ 0 . 8 2 8 ~ 0 . 8 4 8 ~ 0 . 8 1 4 [ 0 . 7 1 0 ~

5 [ 0 . 7 7 ~ 0 . 8 0 8 [ 0 . 8 3 9 ~ 0 . 8 6 1 ] 0 . 8 2 3 ~ 0 . 7 1 0 ~

RAKE NO.

2

1 4 1 6

TUBE NO. I 3 1 4 1 5 1 6 1

0.328 CD,. = 0.150 Bleed

0.210 e x i t s e t t i n g = Closed

n 1 TUBE NO.

1 3 1 4 5 6 io. 838 lo. 867 811 ----

I O . 835 IO. 861 3 .830 0 .728 0.837 0.868 3.808 0.700

0.307 ~~ % = 0.161

Bleed " 0.056 e x i t s e t t i n g = C l o s e d

lr . " TUBE NO. 1

91

Table 3. - TRANSONIC ENGINE-FACE PRESSURE RECOVERY DATA, ptJpt, - Concluded

TUBE NO. 1 I RAm II TUBE. NO.

E lo. 860 10.856

10.868

~~

3 0..872

0.863

0,87 5

_" .

0.87810.868]----

0.874 10.872 10.854)

0.876 10.87630.856 I

TUBE NO. RAKE 11 TUBE NO.

1 .5 10.858

IO. 880

10.858

I ..6 1"" 10.759

I O . 784

NO. 1 2 1 3 2 10.906 10.915

4 10.904 10.906

6 nO.910 10.928

l0.905l0.912

0.909 0.915 10.928 10.934

10.931 10.934

I O . 939 10.938

0.934

0.933 5 ~~0.90810.91710.934

II -- . .

TUBE NO. 1 3 1.4 10.927 10.94k 10.932-[0.947

2

0.908

0.911

=

i - . r.905 10.916

0.901 0.908

0.93110.945b.948L0.915

0.-92510.943 10.94310.900

0.93810:.942 b.943Cl.898

- - ~"

-~ "

0.904

0.904 b.936 .. 10.952 0.918

1.2 a = M, = _____ o.oo - mi/m, = 0.275 CD,. = 0.260 Bleed

Pt2/ptm= 0.906 (x/R)lip= 4.030 A Pt 2 = 0.043 exit setting= Closed

92

Table 4.- INDEX TO FIGURES

F igure

1

2

3 4 5

6-21

22-37

Model photograph

Sketches of model

Design f low f i e l d

Area d i s t r i b u t i o n s

T h e o r e t i c a l mass flow r a t i o

Design Mach number performance

6

7 a 9 10

11

1 2

13 14 15-16 17 18 19 20

21

Superc r i t i ca l pe r fo rmance , b l eed ex i t s e t t i n g A ( a = oo, %p/% = 0 )

I n l e t c o n t r a c t i o n r a t i o

Supe rc r i t i ca l pe r fo rmance ( a = oO, mbp/m, = 0) T o t a l p r e s s u r e d i s t o r t i o n p r o f i l e s ( a = Oo, mbP/m, = 0)

Superc r i t i ca l b l eed b low, i nd iv idua l zones ( a = 00, mbp/m, = 0 )

Bleed lenum chamber pressure recoveries ( a = Oo, mbp P m, = 0)

S t a t i c p r e s s u r e d i s t r i b u t i o n s ( a = 0 , mbP/rq,, = 0)

Flow p r o f i l e s ( a = Oo, mbP/rq,, = 0)

Tolerance t o change i n a n g l e of a t t a c k (qP/m, = 0)

Supe rc r i t i ca l pe r fo rmance wi th bypass (a = 0")

Bypass plenum chamber pressure recoveries ( a = 0')

E f f e c t of bypass on d i s t o r t i o n p r o f i l e s ( a = 0")

S t a t i c p r e s s u r e d i s t r i b u t i o n s w i t h b y p a s s ( a = 0")

Maximum performance a t ang le of a t t a c k w i t h bypass

Superc r i t i ca l pe r fo rmance a t ang le o f a t t ack w i th bypass

0

Off-design Mach number performance

22 Supe rc r i t i ca l pe r fo rmance ( a = oO, mbP/m, = 0)

23 T o t a l p r e s s u r e d i s t o r t i o n p r o f i l e s ( a = Oo, mbP/m, = 0) 24 S u p e r c r i t i c a l b l e e d f l o w , i n d i v i d u a l z o n e s ( a = Oo,

mbp/% = O )

25 Bleed lenum chamber p r e s s u r e r e c o v e r i e s ( a = Oo, q p P m, = 0 )

26-28 S t a t i c p r e s s u r e d i s t r i b u t i o n s ( a = Oo, mbP/h = 0)

29 Flow p r o f i l e s ( a = Oo, mbP/% = 0)

93

Table 4 .- INDEX TO FIGURES - Concluded r ~ . - " . . _ ~ ~ - . . .

Figure

Tolerance t o change in angle of a t t ack (mbP/m, = 0)

31-32 Supercritical performance with bypass (CL = Oo)

33 Bypass plenum chamber pressure recoveries ( a , = 0')

34 Effect of bypass on d i s to r t ion p ro f i l e s (u = 0')

35 Static pressure distributions with bypass ( a = 0")

36 Maximum performance a t angle of attack with bypass

37 Supercritical performance a t angle of a t tack with bypass

38 Maximwn performance a t angle of a t tack (qP/m, = 0)

39 &ximum performance with bypass ( a = 0') 40-43 Effects of uns tar t on the inlet parameters

44-46 Transonic performance . . -. ". - . .~~

94

A-39809

Figure 1.- Model mounted in supe r son ic wind t u n n e l .

-= 0 X R

Cowl bleed e x i t f a i r i n g

Cowl t r a n s l a t i o n

I

Bypass e x i t f a i r i n g

2.825 1 4.180 I \ I rEngine -f ace

r Four s t ru t s equal ly spaced

Centerbody b leed ex i t

f a i r i n g

S ta t ic p ressure rakes (s ix equal ly spaced)

Sleeve t r a n s l a t i o n

I l l I I I '7. I > ' I1

l F i x e d plug

-II/ 4 -

bleed ex i t controls

motor

Lip

( a ) Overall model,

F igwe 2 . - Model.

- = 0.0185 2 R

” - 0.029

Cowl b l e e d e x i t c o n t r o l s ,- Vortex -,- rl

Approx. 0012

S t a t i c p r e s s u r e

Bleed zones

Three rakes ( e q u a l l y s p a c e d )

( b ) I n s t r u m e n t a t i o n , b l e e d a n d b y p a s s c o n f i g u r a t i o n .

F igure 2.- Concluded.

"""" /- ""_ i 0 I 0 2.8 3.2 3 . 6 4.0 4.4

- X

\ R

Figure 3.- Design t h e o r e t i c a l flow f i e l d .

Figure 4. - Area dist r ibut ions.

ID ID

1.0

Figure 5. - I n l e t theoretical mass-flow ratios, Q: = 0".

100

92

.88

.84

.80

Gt* Pt,

76

- 72

.68

.a

H I I l l I I 1 lm

27r VI R k

rI fl I I l l 1 I m m

Design contraction-

ratio

rrrr

Figure 6.- Supercritical performance for various cowl lip positions; & = 3.50, a! = Oo, mbp/& = 0, bleed exit setting A.

101

8.0

0 Cowl l i p p o s i t i o n s for which most of t h e d a t a w e r e o b t a i n e d ,

Maximum p r e s s u r e r e c o v e r y , = 3.50 a = 0"

2.8 3.0 3.2 3 .4 3.6 3.8 4.0 4 .2 4.4

F i g u r e 7.- I n l e t c o n t r a c t i o n r a t i o .

.80

.60

..

103

Bleed exit se t t i ng

A B

C

.8

.7

- r R

.6

-5

.4 &l .9 1.0 d .9 1.0

F 3 - jowl Bottom rake

zu14 LULL

\

I .

0 0 n

.9 1.0 . T I TiwikKd

.9 1.0

BLEED ZONE

1 2 -"

.9

.8

.7

.6

. s .02 .04 .02 .04

3 4 -"" "" mrr ""- +"- I "-"

02 .04 .06

mb 1

0 .02 .04 . c,6 .08

moo

Figure 10. - Supercrit ical bleed flow, individual bleed zones; & = 3.50, ( x / R ) ~ ~ ~ = 2.835, a = O",

mbP/m, = 0.

.80

.40

0

.80

.40

0

1 1 1 1

Bleed exit s e t t i n g

1 4 t

106

70

60

50

40

Theor.

-Centerbody.

20 - el! *

10

0 2.0 k

( pd/pU) centerbody f

p3.03 n /

2.4 2.8 3.2 3.6 4.0 4.4 4.8 5.2 5.6

X - R

A "

z

Theory

6.0 6.4 6.8 7.2

Figure 12. - Static pressure distribution; bleed exit seeting B; m /m, = 0 , = 3-70, (x/R)lip = 2.835, Q: = 0 " .

bP

70

60

50

40

30

20 ”””I_ -lt”-+””-- $

2.0 2.4 2.8 3.2 3.6 4.0 4.4 4.8 5.2 5.6 6.0 6.4 6.8 7.2

70 0 Centerbody

60

50

40

30

Cowl

Cowl l i p Bleed zone

U

20 P

10

I I I I

2.0 2.4 2.8 3.2 3.6 4.0 4.4 4.0 5.2 5.6 6.0 6.4 6.8 7.2

Figure 12. - Concluded.

Cowl rakes .04 - I I

0 0 0 .2 .4 .6 .8 \

1.0 .8 1.0

h R -

.08

.04

0

ptoo

Q Flagged symbols indicate pitot pressures Cowl

Centerbody rakes

x/R = 3.600 Throat rake x/R = 4.200 Centerbody

x/R = 4.400 x/R = 4.225 * -rake

0 .2 .4 .6 .8 .6 .8 1.0 .8 1.0

1.0

-90

.80

-70

.60

.80

1 l A l I I .12 .14 .16 .18 .20

mb 1 m,

Figure 14.- In l e t t o l e rance t o change in apgle of a t tack, b leed exit se t t ing B; M, = 3.50, a = 0.9, mbp/m, = 0 .

111

%n

Ptw

.80

.70

.60

.50

.40

.60

-40

A p t 2

.20

, 2 4 .6 .8 1.0

Figure 15. - Change in mass flow at the engine face for various settings of the bypass exit, bleed exit setting B; M, = 3.50, (x/R)lip = 2.835, a = 0 " .

112

.80

- 70

.60

.50

.60

.40

.20

0

klf f i l l e d symbols,

se t ab 1

1 1 I I

1

Tr m

- .10 .12 .14 .16 .18 .20

mb 1

Figure 16.- Supe rc r i t i ca l performance f o r var ious se t t ings of the bypass e x i t , b l e e d e x i t s e t t i n g B; M, = 3.50, (x/R)lip = 2.835, cy. = Oo.

113

.10 .12 .14 .16 .18 .20

Figure 17.- Bypass plenum chamber pressure recovery, bleed exit setting B; M, = 3.50, a = 0'.

114

Figure 18.- Effect of bypass mass flow on engine-face distortion profiles, maximum pressme recovery, bleed exit setting B; M, = 3.50, Cl = 0'.

4.8 5.0 5.2

0 0 0.01 n 0.02

Figure 19.- Effect of bypass on the static pressure distributions, bleed exit setting B, maximum pressure recovery; M, = 3.50, cx = Oo, (x/R)lip = 2.835.

116

mbl m,

A "

Half f i l l e d symbols ind ica te

SimiI.ar flags indicate the same bypass e x i t opening ,I _ - . - I I . . 1 ~ 1

Figure 20.- Maximum performance a t angle of a t t ack for various amounts of bypass mass flow, b l eed ex i t s e t t i ng B; M, = 3.50.

117

1.00

.80

.60

.40

1.20

. a0

.40

Figure 21.- Superc r i t i ca l performance a t angle of a t tack with and without bypass , b leed exi t se t t ing B; M, = 3.50.

118

.80

Ct* Pto3 .70

.60

- 50

.60

.40

P t 2

.20

.14 .16 .20 .22

(a) M, 5: 3.25, = 2.903

119

.60

.40

.20

0 .12 .14

I

1

I 1 I I

.16 1 .18

mb 1 m,

Bleed exit setting

0 0 n

I I I I I I .

.20

3

A B C

.22

(b) & = 3.00, ( x / R ) ~ ~ ~ = 3.000

Figure 22.- Continued.

120

1.oc

70

.60

.60

.40

A p t 2

.20

0

’ ! I

I. .10 .12 .14 .18 .20

( c > % =’2.75, = 3.230

Figure 22, - Continued.

1 2 1

.12

Bleed ex-: s e t t i n g

0 A O B n C

.14 .16

( d ) = 2.50, ( x / R ) ~ ~ ~ = 3.370

Figure 22. - Continued.

122

1.00

-90

gt2 PtciJ

.80

.70

.60

.60

.40

P t 2

.20

I

I I

I U 1 .06

I 1 I /

Bleed ex i t setting

O A O B n c

. oa .10 .12 .16 mbl m,

( e ) M, = 2.25, ( x / R ) ~ ~ ~ = 3.600

Figure 22.- Continued.

123

1.00

.90

. r 7 0

.60

.60 - -

.40 "

0 .06 . i

I 1 1 I

~ I " .10 .1.2

mb 1

Bleed e x i t setting

A B C

.14 .16 .18

m,

( f ) & = 2.00, ( x / R ) ~ ~ ~ = 3.776

Figure 22.- Continued.

124

.60

.40

Pt*

.20

0 .(

C

.08 .10

Bleed e x i t s e t t i n g

O A O B n c "~ .~

.12 .14 .16

Figure 22. - Continued.

125

I ill

Bleed exit setting

O A O B n c ..

.10 .12 .14 .IC;

(h) Moo = 1.55, = 3.890

Figure 22.- Concluded.

126

Bleed e x i t s e t t i n g

A B C

.8

.7

.6

.5

. b ) / / / / / / f / / / / / / .8 .9 1.0 .8 .9 1.0

Cowl

Bottom rake

.8 .9 1.0 .8 .9 1.0

Figure 23.- O f f design engine-face dis tor t ion prof i les , maximum pressure recovery; a = O o , mbp/m, = 0 .

Bleed ex i t s e t t i n g

O A O B n c

.8

.7

.6

.5

Bottom rake

-4-

-+-

( b ) M, = 3.00

Figure 23.- Continued.

L

.a

.7

.6

- 5

.4 .a .9 1.0

-2Y

.8 .9 1.0

Bottom rake

r u u u 5 -

.8 .g 1.0 .8 .9 1.0

PtB

*tm

( c > M, = 2.75

.a .9 1.0

Bleed e x i t s e t t i n g

O A O B n c

.8 .g 1.0

Figure 23.- Continued.

Bleed ex i t s e t t i n g

O A O B n c

. h r m m r r m n .8 .9 1.0

/mm

Y3- Cowl

Bottom rake

-4Y

C ent erb ody mm hmmm l ? 7 7 7 7 7 h 7 7 T 7

.8 .9 1.0 mTTTTTT7

. 0 .9 1.0

(d) M, = 2.50

Figure 23. - Continued.

.8 .9 1.0 .8 .9 1.0

Bottom rake

Ptm

( e ) M, = 2.25

Figure 23.- Continued.

-5-

.8 .9 1.0

Bleed e x i t s e t t i n g

0 A O B n C

u u u 6 U u L ’

.8 .g 1.0

Bottom rake

Bleed ex i t s e t t i n g

O A O B n c

. 4 m 7 m h T T m ' .a .9 1.0 .

P

Cowl

IT (f) M, = 2.00

Figure 23.- Continued.

-t

.5 - W

.4 7// / / / / / / / / / /7 .8 .g 1.0

Bottom rake

-5 -

W

/ / / / / / / / / / / / / /

.8 .g 1.0

Bleed e x i t s e t t i n g

O A O B n c

-6 LULL’

Figure 23.- Continued.

TOP rake

I

. 4 m i ' .8 .9 1.0

Ptcn

(h) b = 1.55

Figure 23.- Concluded.

Bleed ex i t s e t t i n g

0 A 0 B n c

.8 .9 1.0

I l,,,,J m .8 .9 1.0

1 1 .O”

c”

.6&

3 4 ti7 .06

2 “

I

--I--

LI . ob .06

Bleed zDne

3 r-

4 Bleed es’it

s e t t i n g

”-

O A ”-

02 .04 .06 0 .02 .04 .06

(a) M, = 3.25, (x/R)lip = 2.903

Figure 24.- Off-design supercrit ical bleed f low, individua.1 bleed zones; QI = Oo, mbp/m, = 0.

1

- .08

2

.02 .04 0 .02 .04 “1

(b) = 3.00, ( x / R ) ~ ~ ~ = 3.000

4

”

0 .02 .04

Figure 24.- Continued.

Bleed e x i t

.6 setting O A

n c . u B

-504 .06 .08 1.0

Bleed zone

2 3 4 "- "-

L"

02 .04 c;

.02 .04 0 .02

Figure 24. - Continued.

Bleed zone 1

.5;L"l-"; I I

.02 .04 .06 .08 .10

1

2

.02 .04

3

0 .02 .04

4

" I !

I I ! !

+- I I "I 0 .02

?04 t .06 .08 .10

Bleed zone

2 3 4

t" 0 .02

mb 1 m, -

( e ) M, = 2.25, (x/H)lip =

0 .02

3.600

t-"

3 .02

Figure 24. - Continued. I-

W w

c,

0 P

1.0

.9

.8

1

* 502 I”“

.04 .06 .08

Bleed zone

2

I

“3- ”‘

0 .02

3

0 .02

4

f

0 .02 1 -

( f ) & = 2.00, ( x / R ) ~ ~ ~ = 3.780

Figure 24. - Continued.

1 1.0 ""__

.9

.8

.7 " +"-

Bleed exit .6 setting

O A 0 B A c

.5 .04 .06 .08

Bleed zone

2 3

--p- "d"

I" "

0 .02 .04

mb 1

0 .02

4

I1 1 .02

(g) M, = 1.75, = 3.870

Figure 24. - Continued.

1

Bleed zone

2

mb 1

3

Figure 2b. - Concluded.

0 .02

= 3.890

4

0 .02

.80

.40

0

.80

0

.80

.40

Bleed exit

I I

0 1 .I2

~~ I I I I I I I

.18 .20

(a) M, = 3.25

3 .22

Figure 25.- Off-design bleed plenum chamber pressure recoveries; a = Oo, mbp/mcc = 0.

143

*80 ~i zone

.40

0

.80

.40 E 0

( b ) M, = 3.00

Bleed exil s e t t i n g

"

I

I I--:

I. " .20

I 1 I I

I i 1-

C

B

"

A

2

Figure 25 . - Continued.

144

-80 I -40 I

I

0-

.80 I I

0 I s 4 0 I .10

I 1 1 1 I I I .12 .14 .16 .18 .20 .22

mb 1 m,

~

Figure 25. - Continued.

145

.80

.40

0

.06 .08 .10 .12 .14 .16

1 3 l 4 1 . .

.20

Figure 2 5 . - Continued.

146

-80/

0 I *41

I I 1 1

I

I10 18 .08 .12

mb 1

14 .16

m,

( e ) M, = 2.25

Figure 25 . - Continued.

147

.80

.40

0

I Bleed e x i t

80

40

0 .(

"

.08 .10 .12 .14

r

C

B

~

A

I I I I

1 ~

.18

moo

(f) M, = 2.00

Figure 25.- Continued.

148

.80

.40

0

.80

0

Figure 2 5 . - Continued.

149

111 I1 I111111111111111 I I1 1.1 111.111111 I1

.80

.40

0

. exit , t i n g

1 S .'b 6 .08 .10 .12 .14 .16

Figure 25.- Concluded.

150

IJ Cowl 60 -

! 8

I Cowl lip Bleed zone Engine -face

rake 50 """

40

30

20

10

0 2.0 2.h 2.8 3.2 3.6 4.C 4 .4 4 .8 5 .2 5.6 6.0 6.4 6 . 8 7.2

P

70 0 Centerbody

0 Cowl 60 ! ---L-".~ , ' 1

I I

50

40

30

20

1""" """"", Cowl l i p 'Bleed zone

Engine-face --"

"

"-,

"J

"""

""

10 I""-+" ,-$g """""""- """" 2 " o"l-,+++" ~ --t"------ """

0 """-I ""

2.0 2.4 2.8 3.2 3.6 4.0 4.4 4.8 5.2 5.6 6.0 6.4 6.8 7.2 ""-.

R

( b ) jjt2/ptm = 0.871, rr+,l/mcO = 0.161

Figure 26. - Continued.

P

70 - """"" "

0 Centerbody

c] Cowl -

60 I

1 -~ Cowl l i p Bleed zone Engine-face

rake 50 - I I

40

30

20

10

t 1 I I I

I A I n l @ I " I " I

I

0 2.0 2.4 2.8 3.2 3.6 4.0 4.4 4.3 5.2 5.6 6.0 6.4 6.8 7.2

1

Figure 26. - Concluded.

' Cowl lip I Bleed zone

L" 1 , 1 :

L

"I "c- ~~~"d"""

10 - " 7 8 " fp """"""

+-- - n w 9 "&= . . , " c _ L -7"-

0 - 0 " "- ""- 2.0 2 . k 2.8 3.2 3.6 4.0 4.4 4.8 5.2 5.6 6.0 6.4 6.8 7.2

""""".

X

R

Figure 27.- S t a t i c p r e s s u r e d i s t r i b u t i o n , b l e e d e x i t s e t t i n g B; M, = 2.50, ( X / R ) ~ ~ ~ = 3.370, a = Oo, mbp/m, = 0 .

70 ””

0 Centerbody ”-

Cowl 60 ”

- Cowl l i p Bleed zone

50 - “-

”“”

40 ”-

P - PC0

// /’ Vortex -

generators 30 7- __ ”

“-

20 ”””- - “”””” @”-

1 I

r\ 1 ,

0 o l “ l I ~

2.0 2.4 2.8 3.2 3.6 4.0 4.4 4.8 5.2 5.6 6.0 6.4 6.8 7.2

(b) Pte/pt, = 0.884, mbl/m,, = 0.164

Figure 27. - Continued.

70 . 0 Centerbody

I

I i"J Cowl 60 -

' l l ' l l l l l l ~ l l I l l

y.1" - ~

n "C-+"-J+& 0.oV"""""""""""" 2.0 2.4 2.8 3.2 3.6 4.0 4.4 4.8 5.2 5.6 6.0 6.4 6.8 7.2

c

X

R

Figure 27.- Concluded.

70

60

50

40

30

_i

0 Centerbody

I I [7 Cowl

20 "" - """"""- $-*

10

0 2.0 2.4 2.8 3.2 3.6 4.0 4.4 4.8 5.2 5.6 6.0 6.4 6.8 7.2

( a ) Pt */Ptm = 0.892, mbl/mm = 0.151

Figure 28 . - S ta t ic p ressure d i s t r ibu t ion , b leed ex i t se t t ing B; M, = 2.00, ( X / R ) ~ ~ ~ = 3.780, a = O o , mbp/m, = 0.

70 0 Centerbody

0 Cowl 60

; ! l l ! ' 1

t 1

I r , Cowl l i p Bleed zone ; # I I 1 ;

50

40

30

-"

"

2.0 2.4 2.8 3.2 3.6 4.0 4.4 4.8 5.2 5.6 6.0 6.4 6.8 7.2

(b) Pt2/ptco = 0.865, mbl/m, = 0.140

Figure 28. - Concluded.

Cowl rakes .04 “ I

x/R = 4.448

0 t ? Y - .2 .4 .6 .8 1.0

.8 1.0

(a ) M, = 3.25, (X/R)lip = 2.903

Figure 29.- Off design pitot pressure profiles, maximum pressure recovery, bleed exit setting A; a = oo, q p / m , = 0.

I- o\ 0 Cowl rakes

.04

0 .2 .4 .6 .8 1.0 .8 1.0

Pt JPt, .8 1.0 -

Q Flagged symbols indicate pitot pressures A

Centerbody rakes

x/R = 3.600 x/R = 4.200 04 '

"

0 .e .4 .6 .6 .8 1.0 .6 .8 1.0 .8 1.0

Figure 29. - Continued.

Cowl rakes .04

0 .4 .6 .a 1.0 .a 1.0

Pt JPt, .a 1.0

h R - .08

.04

0 0 .2 .4 .6 .6 .a 1.0

.6 .a 1.0 .a 1.0

ptz Pt,

Figure 29.- Continued.

.04 Cowl rakes

.8 1.0 .8 1.0 Pt Z/P& .a 1.0

Q Flagged symbols i n d i c a t e p i t o t p r e s s u r e s

.08 Centerbody rakes

.04

n \

.2 .4

w 0

I

Y 0

Y

II ct 0

0 0 0

W

IJ P

P ""W Iu

II

w

P 0

vl

"". .

P ""- c

-4 os ul 0 0 0 ""-

II

0

cn ""

0 1

-4 03 \D 0 0 0

"-. 0 0 0

II 0

ul

"- I 0

90

.80

.10

I I I I I I I I L I I# I H I I I I I I I

.12 .14 .16 .18

mbl

moo ( b ) I& = 2.50

Figure 30.- Concluded.

).9" I I I

.20

164

.80

.60

-50

.60

.40

.20

I

1 + .4

4

2 \

b

P 1

c/l 4

.6 .8 1.0

Figure 31.- Off-design change i n mass flow a t the engine face f o r various se t t i ngs of the bypass ex i t , b leed ex i t se t t ing B; Q: = Oo.

165

0 .2 .4 .6 .a 1.0 - m2 moo

(b) M, = 3.00, (X/R)lip = 3.000

Figure 31. - Cont i nued.

166

I

'q

U 0 .2 .a 1.0

167

*pt2

I

I

0 .2 .4 1

.6

0 0 n 0 17 D

0 0.05 0.11 0.20 0.37 0.58

.8 1.0

.9(

.7c

.60

.4c

I 0

f \

0 O n 0 17

y 6

0

0.14 0.26 0.47

0.03

(e) M, = 2.25, ( X / R ) ~ ~ ~ = 3.600

Figure 31.- Continued.

169

m2 m, -

( f ) Mo, = 2.00, ( x / R ) ~ ~ ~ = 3.780

Figure 31. - Continued.

170

I I 1

U 0 .2 .4 .6 .8

m2 mm -

Figure 31.- Continued.

171

0 0 0 0.04

0 .2 .4 .6 .8

Figure 31.- Concluded.

172

I I1 I1 111111 I I 1111

.10

se tab 1

.12 .14 .16 mb 1 m,

. la .20

( a > M, = 3-25> (x /R) l ip = 2.903

Figure 32.- Off-design supercritical performance for var ious s e t t i ngs of t he bypass ex i t , b l eed ex i t s e t t i ng B; CI = Oo.

I 173

.16 .12 .14 .16 .20 mb 1

174

Figure 32. - Continued.

I

1.oc

1 - 9c

'I

ij i

Pt2

-

Pt,

-iO1 .60- I S 6 O r r

~ p t 2 -40r r 0- r .201 .10 12 .14 .16

mb 1

TE -18 .20

Pt,

Pt,

mb 1

3 3 n 3 17 D

0 0.05 0.11 0.20 0.37 0.58

I 1 I I I

.14 .16

176

I

1.0(

.6C

Half filled

0- .06

symbols, ?e table 2

I 1 1 1 I I I ' I I

1 * $ I

1

I I 1

.08 .10 .12 .14 .16

mbl moo

(e) Moo = 2.25, ( X / R ) ~ ~ ~ = 3.600

Figure 32. - Continued.

177

(f) = 2.00, (X/R)~~* = 3.780

Figure 32.- Continued.

178

m 4 0 1

I

0- I .06

Half filled symbols,

see table 2

I I I 1 I I: I

i 0 0 0 0.03 n 0.10 0 0.18 I7 0.31

.08 .10 .12 .14 .16

Figure 32. - Continued.

179

.06 .08 .10 .12

mbl

0 0 0 0.04

I ! 6 .14 .16

180

. l a

Figure 33.- Off-design bypass plenum chamber pressure recovery, bleed exit setting 13; Q: = Oo.

181

.14 .16

mb 1 moo

(b) & = 3.00

b

.18 .20

Figure 33 . - Continued.

182

183

1.0

* 9

.8

.7

.6

-5E, .4 .08 .10 .12 .14

mb 1

I I 1 I

ii I r J

3 0 7 0.05 A 0.11 3 0.20 7 0.37 D -0.58

.16 .18

max

Figure 33. - Continued.

184

.08 .10 .12 .14

( e ) = 2.25

Figure 33.- Continued.

185

.9

.8

-7

.6

- 5

.4 .06 .08 .10 .12

( f ) = 2.00

Figure 33.- Continued.

186

I

l.C

.5

.e

Iitbp - 7

*too

. r

- 5

.4 .06 .08 .10

Figure 33. - Continued.

187

.10 .12

0 0 0 0.04

.14 .16

Figure 33. - .Concluded.

188

Top rake Bottom rake ............................................. ',A\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\

.80 .84 .88

Figure 34.- Off-design effect of bypass on engine-face distortion profiles, maximum pressure recovery, b l e e d e x i t s e t t i n g B; QI = 0 " .

Top rake Bottom rake

Cowl

enterk lody I l I l i I

P t m

( b ) Moo = 3.00

Figure 34. - Continued.

Top rake Bottom rake 1\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\, \\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\

Pt2 Ptw

( c ) M, = 2.75

Figure 34.- Continued.

Bottom rake

4

Figure 34. - Continued.

.a

- 7

. 6

.5

.4 .80 .84 .88 -92 .96 .84

Pt2 Ptco

( e ) M, = 2.25

Figure 34. - Continued.

Top rake Bottom rake

Cowl ...............................

Centerbody

Figure 34. - Continued.

Top rake Bottom rake ...................................... \\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\

Centerbody [7 0.31 .4 \\\<\\\\\!\\\\<\,\\\\!\\\\\I 1 \\\\\\\\\\\\\\\\\\\\\\\\\\\ .88 .92 .96 1.00 .88 .92 .96 1.00

Figure 34. - Continued.

Top rake Bottom rake I \\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\

Cowl

pt,

(h) % = 1 .55

Figure 34. - Concluded.

I

46

44

42

40

44 Centerbody

4*1 4 0 ~

1 38 4.8 5.0 5.2

Engine face

5.4 5.6 5.8

Figure 35.- Off-design effect of bypass on the static pressure distributions, bleed exit setting B, maximum pressure recovery; Q: = Oo.

197

30

28

26

C o w l

"

ff?? kT

Centerbody "~ - - I

4.8 5.0 5.2 5.4

Bypass L

?

. .

.. .

Engine face

5.6 5.8

(b) I%, = 3.00, ( x / R ) ~ ~ ~ = 3.000

Figure 35. - Continued.

198

26

24

22

20

P Po0 -

18

Cowl

-0- a ’r”7

Figure 35.- Continued.

199

14

13

12 5.2 5.4

Centerbody

5.6 5.8

(d) Moo = 2

Figure

Bypass

6.0

= 3.370

0 0 n 0 U D

I

J 1 I

- I I

Engine

0.05 0.11 0.20 0.37 0.58

Face

I I I

6.2

200

I

11

10

9

8 P

Engine face

6.0 6.2 6.4

- X

R

Moo = 2.25, (x/R)lip = 3.600

Figure 35.- Continued.

201

6.2

7 /

7

Bypass

(f) M, = 2.00, ( x / R ) ~ ~ ~ = 3.780

Figure 35.- Continued.

I 4 T I I f I

h g i n e f ace 1

I

g&j

I

I 1 I

7 I 1

I

6.4 6.6

2 02

6.0

P=

I_

C o w l

- Bypass

6.2 6.4 I 1'1

6.6

( g > Moo = 1.75, ( x / R ) ~ ~ ~ = 3.870

Figure 35. - Concluded.

2 0 3

Half filled symbols ind ica te m2/& = 0

.8 1.0

Figure 36.- Off-design maximum performance a t angle of a t t ack for various amounts of bypass mass flow, b l eed ex i t s e t t i ng B.

204

I

mb 1 m,

I - I I -7 1

Half f i l l e d symbols ind ica te m2/% = 0 *-

QI, deg 0 0 0 2

5 0 8

i

0 .2 .4 .6 .8 1.0

mbP - m,

( b ) = 3.00

Figure 36. - Continued.

2 05

mb 1

Figure 36. - Continued.

206

Half f i l l e d symbols ind ica te m 2 / ~ = 0

.601 I l = l " l . o a , 0 deg

0 2 % A 5 0 8

.40 I

0 .2 .4 .6 .a 1.0

mbp m,

( a ) = 2.50

Figure 36. - Continued.

207

" 0 .2 .4 .6

( e ) = 2.25

Figure 36.- Continued.

.8

208

I

.6C

.3c

mb 1 moo

.2c

W 0 .2 .4 .6 .8

(f) = 2.00

Figure 36. - Continued.

2 09

mb 1

i n d i c a t e

0 0 0 2 A 5 0 8

Half f i l l e d symbols mJmw - o

-

0 .2 .4

mbP m,

I 1 I I I I

". I .6 .8

Figure 36. - Concluded.

2 10

I

Half f i l l e d symbols, f ixed bypass e x i t opening

I I 1 - 1 . I L L

I I I

.4 .5 .6 -7 .8 - 9 m2 moa

Figure 37.- Off-design supercr i t ical performance a t angle of at tack with and without bypass, bleed exit sett ing B.

211

1.oc

.8C

.40

.20 7 bypass exit opening

1

.4 .6 .7 .8

m2 moo -

( b ) & = 3.00

Figure 37.- Continued.

.9

212

""

1.0(

.8(

Pt2 Ptcn

.6c

.4c

.2c

.E

Apt* .4

A

I I

T I

Open H a l f

rr I 1 Tr

symbols, I T bypass

f i l l e d symb 01s ,

14 I"

- bypass exit opening

l l r r r

I 1 I 1 1 1 .4 I I

-5 .6 .7 .8 m2 moo -

Figure 37. - Continued.

213

pt,

1.00

80

.60

.20

.2 .3 .4 .5 .6 .7

Figure 37. - Continued.

2 14

I

.8C

.6c

.4c

1.2

.E

C .1

I r 7 mbp = 0.14

1 Open symbols, bypass = 0 Half f i l l e d symbols, f ixed

Trrl"H bypass e x i t opening

.2

( x / ~ ) l i p 3.600 3.624 3.666 3.704

03 .4 m2

.6

m,

( e ) M, = 2.25

Figure 37. - Continued.

215

1

%2

PtaJ

. . 00

.80

.60

Open symbois, bypass Half filled symbols, bypass exit opening .40

1

Apt2

.2

.a

.4 1

0 0 .1

Figure 37.- Continued.

216

I I

4., I 11 .4 - 5

1.00

.8c - Pt2

Ptco

.60

.40

.8

.4

0 .1 .e - 3

3 moo

Figure 37. - Concluded.

217

1.00

.80

.60

.40

mb 1 m, - .20

.10

.8

.4

1.2 1.6 2.0 2.4 2.8 3.2 3.6 rJE,

( a ) Bleed e x i t s e t t i n g A

Figure 38.- Maximum performance a t angle of a t tack ; mbP/rn, = 0.

218

I

1.0(

.8(

.6(

.4(

0 O n 0

.a

.4

0 I .2

e A

0 2 5 8

' 1

1.6 2.0 2.4 2 .a 3.6

Moo

( b ) Bleed e x i t s e t t i n g B

Figure 38. - Continued.

219

*b 1 m,

.20

.10

0

1.2 1.6 2.0 2.4 2.8 3.6

Moo

( c ) Bleed e x i t s e t t i n g C

Figure 38.- Concluded.

220

I

(%) 'tw max

.3O

.20

Half f i l l e d symbols ind ica te m 2 / m , = 0

0 .2 .4 .6 mbP - m,

( a ) Bleed e x i t s e t t i n g A

.8 1.0

Figure 39.- Maximum performance for var ious amounts of bypass mass flow; a = oo.

2 2 1

222

Half f i l l e d symbols ind ica te m2/% = 0

0 .2 .4 .6

mbP moo

(b) Bleed exit setting B

Figure 39.- Continued.

- 8 1.0

1.00

i I

.80

.20

0

Half f i l l e d symbols indicate mz/moo = 0 I

0 .2 .4 .6 .8 1.0

mbP moo

( c ) Bleed e x i t s e t t i n g C

Figure 39. - Concluded.

2 23

N N P

0 I n l e t u n s t a r t s 0 I n l e t restarts

1””-

3 . 4 L F ” ” I

I ‘7 ””- ”””-

3.2”””- ””-7

,-,”””. ””7

3.0”””” ?-

-,””””. T-” Moo

Figure 40.- I n l e t unstart and r e s t a r t cowl-lip pos i t i on ; a = 0..

I

.80

- Pt2

ptw .40

0

.20

mb 1 m,

.10

0

.4(

.2(

7="k H I ! I !

0 Maximum pressure @ In le t uns ta r ted

1 1 . 1 I

Q 7 \

1.6 2.0

1 7 1 1 1 1 4 T

2.4 2.8 3 .2 3.6

Moo

225

mb 1 m,

0 Maximum pressure recovery I n l e t unstarted Bleed zone 1 4

.04

I I 1 1 1 1 1

.04

0

.04

0

.12

.08

.04

0 1.2 1.6 2.0 2.4

Moo

1 T 1 1 1

1 2.8 3 - 2 3.6

Figure 42.- Effect of uns ta r t ing t he i n l e t on the individual bleed f lows, b l e e d e x i t s e t t i n g B; a = O o , mbp/m, = 0.

226

I

1.2 1.6 2.0 2.4 2.8 3.2 3.6

M,

Figure 43.- Effect of uns ta r t ing the in le t on the individual bleed plenum chamber pressure recoveries, bleed exit sett ing B; a! = O", mbp/m, = 0.

227

2 2 8

.4

c D a .2

0 .16 .20 .24 .28 .32 .36

mi m,

~

( a ) M, = 0.6

Figure 44.- Transonic performance, bleed e x i t s e t t i n g B; a = 00, mbp/m, = 0.

I

.E

Apt* -4

0

.4

.2

IT I I I 1 1 1

0 .16 .20 .24 .28 .32 .36

mi m, -

Figure 44.- Continued.

229

.4

.2

0 .16 .20 .24 .28 - 32 .36

- mi moo

( c ) I&, = 0.8

Figure 44. - Continued.

230

I

0

.4

.2 %

0 .16 .20 .24 .2a .32 .36

mi mm -

Figure 44.--Continued.

231

.4

.2

n

I I I I ! I ! I 1 I I I "

.16 .20 .24 .28 .36

m i m,

( e ) M, = 1.0

Figure 44.- Continued.

232

1.00

-90

.60

.a

Apt2 .4

0

.4

.2

n

L

1 I 1 1 i

I I I A=

.16 .20 .24 .2a - 32 .36 U

(f) = 1.1

Figure 44. - Continued.

233

.4

.2

n

.16 .20 .24 .28 .32 .36 u

mi nbo -

Figure 44. - Continued.

2 34

9(

Pt* .8(

Ptm

.7(

.6(

.16 .20

I I I I

I B

.24 .28 - 32 .36 mi - moo

Figure 44.- Concluded.

235

%,

.6 Open symbols, bleed exits opened Half f i l l e d symbols, b leed ex i t s c losed I

.16 .20 .24 .28

(a) !& = 0.6

I 1 .32 .36

Figure 45. - Transonic performance with and without bleed; a = 0" , mbp/m, = 0.

2 36

0 4.180

3.880 0 4.030

Open symbols, bleed exits opened H a l f f i l l e d symbols, b l eed ex i t s c lo se t

0 .16 .24 .28

mi - moo

S

* 32 .36

(b) I& = 0.8

Figure 45. - Continued.

237

1

Open symbols, b leed ex i t s opened Half f i l l e d symbols, b leed ex i t s

.24

I I I I

f .28 .32

( c ) M, = 1.0

Figure 45 . - Continued.

.36

1. oc

.go

Gt2

PtcxJ .80

* 70

.60

.a

.6

Apt2 .4

.2

0 .16

Open symbols, bleed exits opened Half filled symbols, b l eed ex i t s closed

I I - 1 -

.20

( a ) I& = 1.2

Figure 45.- Concluded.

239

I I 11111 I

1.00

90

Open symbols, b leed ex i t s opened Half f i l l e d symbols, bleed ex i t s closec

.80

.e

.4

.2

0 .16 .20 .24

( a ) M, = 0.6

I 1 I ! :

I .28 .32 -36

240

1.00

.80

.6

.4

Apt2

.2

0 L

Open symbols, bleed exi ts opened Half f i l l e d symbols , bleed e x i t s c l o s e d

i

0 0 n 0

.16 .20

T T T I $ .24

I T Tr Tr g .28

( b ) = 0.8

.32 .36

Figure 46. - Continued.

24 1

242

Open symbols, b leed ex i t s opened Half f i l l e d symbols, bleed ex i t s c losed

.70 "

. I

.2

0 .16 .20 .24 .2a .32 .36

mi - moo

Figure 46.- Concluded.

NASA-Langley, 1970 - 28 A-3516

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